Calorific value

ABSTRACT

A method of checking refuelling of an aircraft comprising a gas turbine engine and a fuel tank arranged to provide fuel to the gas turbine engine comprises: receiving an input of calorific value data for fuel provided to the aircraft on refuelling; independently determining at least one of: (i) the calorific value of fuel supplied to the gas turbine engine in use; and (ii) the calorific value of the fuel provided to the aircraft on refuelling; and providing an alert if the determined calorific value is inconsistent with the calorific value data input received

CROSS-REFERENCE TO RELATED APPLICATIONS

This specification is based upon and claims the benefit of priority fromUK Patent Application Number 2205344.1 filed on 12 Apr. 2022, the entirecontents of which are incorporated herein by reference.

BACKGROUND Field of the Disclosure

The present disclosure relates to aircraft propulsion systems, and tomethods of operating aircraft involving the management of fuels ofdifferent types, including control of refuelling, and to methods ofmodifying aircraft so as to allow such operating methods to beimplemented. The present disclosure further relates to methods of, andsystems for, determining one or more fuel characteristics of an aviationfuel. The present disclosure further relates to operating an aircraftaccording to the determined fuel characteristics.

Description of the Related Art

There is an expectation in the aviation industry of a trend towards theuse of fuels different from the traditional kerosene-based jet fuelsgenerally used at present.

SUMMARY

According to a first aspect there is provided a method of refuelling anaircraft comprising a gas turbine engine and a fuel tank arranged toprovide fuel to the gas turbine engine, the method comprising:

-   -   obtaining an amount of energy required for an intended flight        profile;    -   obtaining a calorific value of fuel available to the aircraft        for refuelling;    -   calculating the amount of the available fuel needed to provide        the required energy; and    -   refuelling with the calculated amount of the available fuel.

Knowledge of the calorific value(s) of fuel(s) available to an aircraftcan allow more efficient, tailored, control of the propulsion system.For example, changing to a fuel with a higher calorific value may allowa smaller amount (mass or volume) of fuel to supply an aircraft's energyneeds for a flight. In addition, as more power is needed to lift agreater mass of fuel, taking sufficient fuel for the intended flight(including a safety margin above the expected energy demand), but notcompletely filling the tank(s), may provide an efficiency bonus byreducing take-off weight of the aircraft. Knowledge of the calorificvalue of the fuel can therefore be used as a tool to improve aircraftperformance, e.g. avoiding carrying excess fuel weight.

The calculating the amount of the available fuel needed to provide therequired energy may comprise obtaining an energy content of fuel alreadyin the fuel tank(s)—e.g. that remaining from a previous refuellingevent - and subtracting that from the determined amount of energyrequired for the intended flight profile.

The obtaining the calorific value of the fuel available to the aircraftfor refuelling may comprise receiving an input of calorific value datafor the available fuel, e.g. via a user interface.

The obtaining the calorific value of the fuel available to the aircraftmay comprise receiving calorific value data in an electroniccommunication, e.g. from the reading of a barcode or QR code associatedwith the fuel supply, or receiving a message from the fuel supplier orrefuelling facility.

The obtaining the calorific value of the fuel available to the aircraftmay comprise chemically and/or physically determining the calorificvalue of the available fuel, optionally by performing one or more of:

-   -   i. identifying a tracer in the available fuel, such as a dye, or        a trace substance used as a marker, and looking up a calorific        value corresponding to that tracer (a trace substance inherently        present in the fuel which may vary between fuels may be used to        identify a fuel, and/or a substance may be added deliberately to        act as a tracer);    -   ii. inferring the calorific value from one or more detected        physical or chemical properties of the available fuel; and/or    -   iii. combusting a sample of the available fuel to determine its        calorific value directly, optionally using the gas turbine        engine.

The chemically and/or physically determining the calorific value of theavailable fuel may be performed onboard the aircraft, and optionally maybe performed in operation of the aircraft, e.g. powering lighting,heating, and/or air conditioning whilst stationary/before refuelling iscomplete.

Especially (but not only) in scenarios in which the calorific value ismanually input, the method may further comprise performing a check toverify the calorific value/input data, the check comprising measuringthe calorific value of the fuel in use in the gas turbine engine,optionally during operation of the aircraft prior to take-off (e.g.taxiing, or at-gate operations), and comparing that to the valueobtained by a different route. In the event of a mis-match beyond athreshold, the aircraft may be returned to the gate/refuelling mayrecommence.

The determining the calorific value of the fuel may be performed bymonitoring engine parameters during a first time period of aircraftoperation during which the gas turbine engine uses the fuel; anddetermining the calorific value of the fuel based on the monitoredengine parameters. This determination step may be used to determine thecalorific value of the fuel already in the aircraft's tanks—e.g. in aflight prior to the refuelling event—such that the energy remaining inthe tanks can be calculated and deducted from the amount needed onrefuelling. This determination step may also be used for the new fuel tobe provided to the aircraft—e.g. combusting a small sample in the/a gasturbine engine whilst the aircraft is stationary and before refuellingis complete, so as to determine how much more fuel to request or accept,or as part of checks to confirm the calorific value obtained onrefuelling.

The gas turbine engine may be a main, propulsive, gas turbine engine, ormay be a gas turbine engine of an auxiliary power unit (APU), which mayor may not be arranged to provide propulsive power to the aircraft. Insome cases, fuel properties may be determined by combusting a smallsample in the APU before starting the main engine(s). In additional oralternative examples, the combustion of a fuel sample for determiningfuel characteristics may be done in one or more of the main propulsionengine(s).

According to a second aspect, there is provided a propulsion system foran aircraft, the propulsion system comprising:

-   -   a gas turbine engine;    -   a fuel tank arranged to contain a fuel to power the gas turbine        engine; and    -   a refuelling manager arranged to:        -   obtain an amount of energy required for an intended flight            profile;        -   obtain a calorific value of fuel available to the aircraft;        -   calculate the mass or volume of the available fuel needed to            provide the required energy; and        -   output the mass or volume of the available fuel needed so as            to allow the aircraft to be refuelled accordingly.

The amount of energy required for an intended flight profile may includea safety margin beyond the amount expected to be needed for the flight.The size of the safety margin may be decided based on a variety offlight, environmental condition, and aircraft parameters.

The propulsion system may further comprise one or more sensors arrangedto physically and/or chemically detect one or more properties orcharacteristics of the fuel. The one or more fuel properties/the sensordata may then be used to infer or calculate the calorific value of thefuel—this may be used for either or both of the fuel already present inthe aircraft's tank(s) prior to refuelling, and the fuel being providedto the aircraft on refuelling. Sensor type and location may be selectedaccordingly.

The propulsion system may be arranged to implement the method of thefirst aspect.

The gas turbine engine optionally comprises:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor; and    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades and being arranged to be driven by an        output from the core shaft.

According to a further aspect, there is provided a non-transitorycomputer readable medium having stored thereon instructions that, whenexecuted by a processor, cause the processor to:

-   -   obtain an amount of energy required for an intended flight        profile of an aircraft;    -   obtain a calorific value of fuel available to the aircraft;    -   calculate the amount of the available fuel needed to provide the        required energy; and    -   output the mass or volume of the available fuel needed so as to        allow the aircraft to be refuelled accordingly.

The instructions may be further arranged to cause the processor tocontrol fuel input to the aircraft, such that the aircraft is refuelledwith the calculated amount of the available fuel.

The instructions may be arranged to cause the processor to perform themethod of the first aspect.

According to a third aspect, there is provided a method of checkingrefuelling of an aircraft comprising a gas turbine engine and a fueltank arranged to provide fuel to the gas turbine engine, the methodcomprising:

-   -   receiving an input of calorific value data for fuel provided to        the aircraft on refuelling;    -   independently determining at least one of:        -   (i) the calorific value of fuel supplied to the gas turbine            engine in use; and        -   (ii) the calorific value of the fuel provided to the            aircraft on refuelling; and    -   providing an alert if the determined calorific value of fuel is        inconsistent with the calorific value data input received.

As used above, “independently” determining the calorific value meansdetermining the calorific value with no use of, or reference to, theinput calorific value data—two values for the calorific value cantherefore be obtained separately and compared. The two values maytherefore be obtained in completely different ways—e.g. one may beprovided by a fuel supplier, and one may be calculated from sensor data.

The inventors appreciated that, as different fuels can have differentproperties, whilst still conforming to the standards, knowledge of thefuel(s) available to an aircraft can allow more efficient, tailored,control of the propulsion system. Verification systems, for exampleimplementing a check of fuel properties if changes to aircraft controlare to be made based on the fuel characteristics, may therefore beimplemented to build trust in, and improve reliability of, such newpropulsion system control techniques.

The checking of the third aspect may be performed for safety reasons—toensure that the total energy in fuel onboard the aircraft is sufficientfor the intended flight, including any safety margin provided (e.g. incase of adverse weather or a need to divert to a different airport).

The checking of the third aspect may be performed for aircraftperformance optimisation reasons—confidence in knowledge of the fuel(s)onboard the aircraft may allow engine operation to be tailored to theavailable fuel(s).

The method may be performed by an aircraft system, and/or by an off-wingunit.

The gas turbine engine may be a main, propulsive, gas turbine engine ora gas turbine engine of an auxiliary power unit (APU), which may or maynot be arranged to provide propulsive power to the aircraft.

The method may comprise determining the calorific value of the fuelprovided to the aircraft on refuelling in a fuel testing unit. The fueltesting unit may be provided off-wing at a refuelling site.

The determining the calorific value of the fuel supplied to the gasturbine engine in use may be performed during at least one of: (i) oneor more operations on the ground prior to take-off (e.g. engine warm-upand/or taxiing of the aircraft), and (ii) climb. It will be appreciatedthat performing the determination relatively early in the flight mayfacilitate taking appropriate corrective action if the determined valueis inconsistent with the calorific value data input received. Ideally,the check may be performed before the aircraft leaves the ground, incase the fuel is not as expected to such an extent that refuelling isadvisable.

The determining the calorific value of the fuel supplied to the gasturbine engine in use may be performed by burning fuel taken from thefuel tank arranged to be used to supply fuel to the gas turbine enginein an auxiliary power unit (APU) of the aircraft and determining thecalorific value.

The determining the calorific value of the fuel supplied to the gasturbine engine in use may be performed by sensing engine parameters (ofthe propulsive gas turbine engine and/or of a gas turbine engine of anAPU) during a first time period of aircraft operation during which thegas turbine engine uses the fuel; and determining the calorific value ofthe fuel based on the monitored engine parameters. Optionally, the oneor more engine parameters may be monitored over the first time period.

The determining the calorific value of the fuel provided to the aircrafton refuelling may be performed by identifying a tracer in the fuelprovided to the aircraft (e.g. a dye or a marker trace element), andlooking up a calorific value corresponding to that tracer.

The determining the calorific value of the fuel provided to the aircrafton refuelling may be performed by inferring the calorific value from oneor more detected physical or chemical properties of the available fuel.

The receiving an input of calorific value data for fuel provided to theaircraft on refuelling may comprise receiving data input via a userinterface of the aircraft, e.g. a typed-in value or value selected froma menu.

The receiving an input of calorific value data for fuel provided to theaircraft on refuelling may comprise receiving data electronicallycommunicated to the aircraft, e.g. sent by a supplier or obtained fromreading a bar code, QR code, or other code associated with the providedfuel.

According to a fourth aspect, there is provided a propulsion system foran aircraft, the propulsion system comprising:

-   -   a gas turbine engine;    -   a fuel tank arranged to contain a fuel to power the gas turbine        engine; and    -   a fuel tracking system arranged to:        -   receive an input of calorific value data for fuel provided            to the aircraft on refuelling;    -   determine at least one of:        -   (i) the calorific value of fuel supplied to the gas turbine            engine in use; and        -   (ii) the calorific value of the fuel provided to the            aircraft on refuelling; and    -   provide an alert if the determined calorific value of the fuel        is inconsistent with the calorific value data input received.

The gas turbine engine may comprise:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor; and    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades and being arranged to be driven by an        output from the core shaft.

The propulsion system may further comprise one or more sensors arrangedto physically and/or chemically detect one or more properties of thefuel, the one or more fuel properties being used to infer or calculatethe calorific value of the fuel. These sensors may be used indetermining the calorific value of either or both of the fuel suppliedto the gas turbine engine in use, and the fuel being provided to theaircraft on refuelling. Sensor type and location may be selectedaccordingly.

The propulsion system may further comprise one or more sensors arrangedto sense at least one engine parameter during a first time period ofaircraft operation during which the gas turbine engine (a propulsive gasturbine engine and/or a gas turbine engine of an APU) uses the fuel; theat least one engine parameter being used to determine the calorificvalue of the fuel.

The propulsion system may be arranged to implement the method of thethird aspect.

According to a fifth aspect, there is provided a method of refuelling anaircraft comprising a gas turbine engine and a fuel tank arranged toprovide fuel to the gas turbine engine, the method comprising:

-   -   determining an amount of energy required for an intended flight        profile (optionally including a safety margin as mentioned        above);    -   receiving an input of calorific value data for the available        fuel;    -   calculating the amount of the available fuel needed to provide        the required energy;    -   outputting the calculated amount of the available fuel needed so        as to allow the aircraft to be refuelled accordingly;    -   determining at least one of:        -   (i) the calorific value of fuel supplied to the gas turbine            engine in use after refuelling; and        -   (ii) the calorific value of the fuel provided to the            aircraft on refuelling; and providing an alert if the            determined calorific value of fuel is inconsistent with the            calorific value data input received.

The calculating the amount of the available fuel (e.g. from a refuellingvessel or pipeline) needed to provide the required energy may compriseobtaining a calorific value of fuel already in the fuel tank, and usingthat to calculate the total energy available from fuel already onboard,optionally using one or more calorific values of the one or more fuelsonboard and the amount of each. That value may then be subtracted fromthe determined amount of energy required for the intended flightprofile. The calculating the amount of the available fuel needed toprovide the required energy may therefore comprise obtaining an energyvalue of fuel already in the fuel tank, and subtracting that from thedetermined amount of energy required for the intended flight profile.

The method may further comprise calculating a calorific value for themixed fuel after refuelling. The calorific value of fuel supplied to thegas turbine engine in use may therefore be determined to be inconsistentwith the calorific value data input received only when the calculatedcalorific value for the mixed fuel does not match the determinedcalorific value of fuel supplied to the gas turbine engine in use.

According to a sixth aspect, there is provided a method of determining acalorific value of fuel supplied to a gas turbine of an aircraft, themethod comprising:

-   -   sensing one or more engine parameters during a first time period        of aircraft operation during which the gas turbine engine uses        the fuel; and    -   based on the one or more sensed engine parameters, determining a        calorific value of the fuel.

The sensing may be repeated such that the at least one engine parameteris monitored over time, and/or to provide a check of a first sensedvalue.

As different fuels can have different properties, such as differentcalorific values, whilst still conforming to the standards, the same gasturbine engine may perform differently when provided with a differentfuel. The performance of the gas turbine engine itself can therefore beused to determine calorific value of a fuel currently being provided toa gas turbine engine. Knowledge of the fuel(s) available to an aircraftis beneficial as it can allow more efficient, tailored, control of thepropulsion system, and/or tailored fuelling.

The first time period of aircraft operation may comprise at least one of(i) a period of engine operation prior to take-off, and (ii) a period oftime spent climbing.

The first time period of aircraft operation may comprise a period oftime spent warming up the engine prior to any movement, a period of timetaxiing prior to take-off, and/or a period of time spent climbing. Inother examples, the first time period may be during cruise. It will beappreciated that performing the determination relatively early in theflight may facilitate making use of the calorific value determined. Thedetermination may be performed before the aircraft leaves the ground, incase the fuel is not as expected to such an extent that refuelling isadvisable, or so as to judge how much of a fuel to load. In otherimplementations, the calorific value may be used to influence propulsionsystem control in flight/for the remainder of the flight following thedetermination.

The sensed/monitored engine parameters may include one or more of fuelflow rate, shaft speed, combustor temperature rise, thrust generated,and exhaust temperature.

The method may further comprise performing a check to verify thedetermined calorific value, the check comprising comparing the calorificvalue of the fuel determined from the engine parameters to a calorificvalue of the fuel which is:

-   -   a) supplied to the aircraft (e.g. by an electronic        communication, or using a graphical user interface);    -   b) calculated from detected fuel properties (e.g. from sensor        data); and/or    -   c) determined from one or more engine parameters sensed in a        different time period of aircraft operation.

The calorific value of the fuel may be determined as a relative value,or change, compared to that of a different fuel used by the same gasturbine engine. A change in the monitored/sensed engine parameter(s) onchanging from one fuel to the other may be used to determine the changein calorific value.

One or more engine parameters may therefore be sensed/monitored in twodifferent time periods—one each for the two different fuels. The changein fuel may be the only change made to engine control between the twotime periods. The two time periods may also be selected such thataltitude and/or other external parameters are at least substantially thesame for both, and may therefore be selected to be close to each otherin time, if not immediately consecutive. An interval may be left betweenthe two time periods to allow for any transient behaviour around thechange in fuel.

The engine parameters may instead be monitored over a single time periodduring which the change between the two different fuels is made. Thechange in fuel may be the only change made to engine control during thetime period. Any transient behaviour around the change in fuel may beused in determining the calorific value in some examples.

The engine parameters may be or comprise shaft speed and fuel flow rate.

The engine parameters may be or comprise temperature rise across thecombustor and fuel flow rate.

The gas turbine engine which is monitored to determine calorific valueof the fuel being combusted may be a propulsive gas turbine engine ofthe aircraft.

The gas turbine engine which is monitored to determine calorific valueof the fuel being combusted may be a gas turbine engine of an auxiliarypower unit of the aircraft. In such cases, the calorific value may bedetermined prior to starting a propulsive gas turbine engine of theaircraft.

According to a seventh aspect, there is provided a method of operatingan aircraft comprising a gas turbine engine and a fuel tank arranged tosupply fuel to the gas turbine engine, the method comprising:

-   -   sensing one or more engine parameters during a first time period        of aircraft operation during which the gas turbine engine uses        the fuel;    -   based on the sensed engine parameter(s), determining data        relating to or comprising a calorific value of the fuel; and    -   adjusting (e.g. optimising) engine performance during a second        time period of aircraft operation based on the determined data.

The sensing may be repeated such that the at least one engine parameteris monitored over time, and/or to provide a check of a first sensedvalue.

Knowledge of the calorific value of the fuel can therefore be used as atool to improve aircraft performance in flight, e.g. gauging thelikelihood of forming contrails behind an aircraft in given atmosphericconditions and changing fuel source or altitude mid-flight asappropriate, or using a lower calorific value fuel for lower thrustoperations.

The first time period of aircraft operation may comprise at least one of(i) a period of engine operation prior to take-off, and (ii) time spentclimbing. The first time period of aircraft operation may comprise atleast one of time spent warming up the engine prior to movement, timespent taxiing prior to take-off, and time spent climbing. Alternativelyor additionally, the first time period may be early on within a cruiseperiod (e.g. within the first 1%, 5% or 10% of the cruise time), soallowing for the optimisation of engine performance thereafter.

The second time period of aircraft operation may be a period of timespent at cruise, and optionally may be the entirety of cruise.

The adjusting engine performance may comprise at least one of:

-   -   changing fuel flow rate;    -   changing pump spill;    -   changing altitude;    -   changing guide vane staging; and    -   changing fuel.

According to an eighth aspect, there is provided a power system for anaircraft, the power system comprising:

-   -   a gas turbine engine;    -   a fuel tank arranged to contain a fuel to power the gas turbine        engine; and        -   a fuel tracking system arranged to:            -   sense one or more engine parameters during a first time                period of aircraft operation during which the gas                turbine engine uses the fuel; and            -   determine a calorific value of the fuel based on the                sensed engine parameter(s).

The gas turbine engine may comprise an engine core comprising a turbine,a compressor, and a core shaft connecting the turbine to the compressor;and a fan located upstream of the engine core, the fan comprising aplurality of fan blades and being arranged to be driven by an outputfrom the core shaft.

The gas turbine engine may be arranged to provide propulsive power tothe aircraft. The power system may therefore be described as apropulsion system.

Alternatively, the gas turbine engine may be an Auxiliary Power Unit(APU) of the aircraft and may not be arranged to provide propulsivepower.

The power system may further comprise at least one sensor arranged tosense the at least one engine parameter.

The power system may further comprise one or more sensors arranged tophysically and/or chemically detect one or more properties orcharacteristics of the fuel, the one or more fuel properties orcharacteristics being used to infer or calculate the calorific value ofthe fuel.

According to a ninth aspect, there is provided a propulsion system foran aircraft, the propulsion system comprising a gas turbine engine andone or more fuel tanks arranged to contain fuel to power the gas turbineengine, one or more of the tanks containing a fuel which is asustainable aviation fuel—SAF—or is a fuel blend including SAF, the SAFhaving a density of between 90% and 98% of the density, PK, of keroseneand a calorific value of between 101% and 105% the calorific value, CVK,of kerosene.

The propulsion system therefore comprises at least one fuel tankarranged to store fuel to power the gas turbine engine, wherein thestored fuel comprises at least a proportion of a sustainable aviationfuel—SAF. If the stored fuel is a SAF-blend as opposed to purely SAF,the blend's calorific value and density will lie between those of theSAF and those for kerosene.

The gas turbine engine comprises:

-   -   a combustor; and    -   a fuel pump arranged to supply a fuel from one or more of the        fuel tanks to the combustor at an energy flow rate, C, the fuel        pump having an inlet arranged to receive fuel from one or more        of the fuel tank(s) and an outlet arranged to output the fuel at        a pump output volumetric flow rate, Q, a proportion of the        output fuel being provided to the combustor and the remainder        being recirculated (i.e. sent from the pump outlet back to the        inlet, either directly or indirectly via one or more systems or        components), the percentage of fuel passing through the pump        which is recirculated/which is not provided to the combustor        being referred to as a spill percentage, and wherein the fuel        supplied to the pump comprises X % SAF, where X % is in the        range from 5% to 100%, with any remainder of the fuel being        kerosene, and has a density, ρ_(F), and a calorific value        CV_(F).

The propulsion system is arranged such that:

-   -   the (dimensionless) fuel-change spill ratio, R_(s), of:

$R_{s} = \frac{{spill}{percentage}{at}{cruise}{using}{kerosene}}{{spill}{percentage}{at}{cruise}{using}{the}{fuel}{with}X\%{SAF}}$

-   -   is equal to:

$\frac{Q - \left( {C/\left( {{CV}_{K} \times \rho_{K}} \right)} \right)}{Q - \left( {C/\left( {{CV}_{F} \times \rho_{F}} \right)} \right)}.$

X may be greater than 50, such that the fuel supplied to the combustoris more than 50% SAF.

It will be appreciated that different tanks may hold fuels with adifferent % SAF.

The gas turbine engine may be arranged such that, for an engine with amaximum take-off thrust in the range from 400 kN to 500 kN, R_(s) is inthe following range at cruise:

${\frac{Q - \left( {\left( {Q - 5.8} \right)/\left( {0.014 \times {CV}_{K} \times \rho_{K}} \right)} \right)}{Q - \left( {\left( {Q - 6.31} \right)/\left( {0.014 \times {CV}_{F} \times \rho_{F}} \right)} \right)} \leq R_{s} \leq \frac{Q - \left( {\left( {Q - 6.31} \right)/\left( {0.014 \times {CV}_{K} \times \rho_{K}} \right)} \right)}{Q - \left( {\left( {Q - 5.8} \right)/\left( {0.014 \times {CV}_{F} \times \rho_{F}} \right)} \right)}},$

and optionally in the range:

$\frac{Q - \left( {\left( {Q - 5.8} \right)/\left( {0.014 \times {CV}_{K} \times \rho_{K}} \right)} \right)}{Q - \left( {\left( {Q - 5.8} \right)/\left( {0.014 \times {CV}_{F} \times \rho_{F}} \right)} \right)} \leq R_{s} \leq \frac{Q - \left( {\left( {Q - 6.} \right)/\left( {0.014 \times {CV}_{K} \times \rho_{K}} \right)} \right)}{Q - \left( {\left( {Q - 5.8} \right)/\left( {0.014 \times {CV}_{F} \times \rho_{F}} \right)} \right)}$

where Q (fuel flow rate) is measured in litres per second, CV (calorificvalue) in MJ/kg, and ρ (density) in kg per litre, with the K and Fsubscripts used for kerosene and the SAF or SAF-blend fuel,respectively.

The gas turbine engine may be arranged such that, for an engine with amaximum take-off thrust in the range from 300 kN to 350 kN, R_(s) is inthe following range at cruise:

$\frac{Q - \left( {\left( {Q - 4.37} \right)/\left( {0.027 \times {CV}_{K} \times \rho_{K}} \right)} \right)}{Q - \left( {\left( {Q - 4.87} \right)/\left( {0.027 \times {CV}_{F} \times \rho_{F}} \right)} \right)} \leq R_{s} \leq \frac{Q - \left( {\left( {Q - 4.87} \right)/\left( {0.027 \times {CV}_{K} \times \rho_{K}} \right)} \right)}{Q - \left( {\left( {Q - 4.37} \right)/\left( {0.027 \times {CV}_{F} \times \rho_{F}} \right)} \right)}$

and optionally in the range:

${\frac{Q - \left( {\left( {Q - 4.37} \right)/\left( {0.027 \times {CV}_{K} \times \rho_{K}} \right)} \right)}{Q - \left( {\left( {Q - 4.87} \right)/\left( {0.027 \times {CV}_{F} \times \rho_{F}} \right)} \right)} \leq R_{s} \leq \frac{Q - \left( {\left( {Q - 4.47} \right)/\left( {0.027 \times {CV}_{K} \times \rho_{K}} \right)} \right)}{Q - \left( {\left( {Q - 4.37} \right)/\left( {0.027 \times {CV}_{F} \times \rho_{F}} \right)} \right)}},$

where, as above, Q is measured in litres per second, CV in MJ/kg, and pin kg per litre.

The SAF proportion (X%) may be gravimetric.

The gas turbine engine may be arranged such that R_(s)≤1.04.

The gas turbine engine may be arranged such that R_(s)≥1.003.

The gas turbine engine may be arranged such that R_(s)≥1.014. It may bethat R_(s)≥1.014 when the SAF proportion is greater than or equal to75%.

In addition to considering fuel properties, R_(s) may be varied based onone or more of ambient temperature, altitude, and stage of cruise.

The gas turbine engine may be arranged such that R_(s) decreases by lessthan 0.15% between the beginning and end of cruise, for a constanttemperature and altitude.

The gas turbine engine may be arranged such that R_(s) decreases by atleast 0.11% when altitude increases by at least 600 m.

The gas turbine engine may comprise an engine core comprising a turbine,a combustor, a compressor, and a core shaft connecting the turbine tothe compressor; and a fan located upstream of the engine core, the fancomprising a plurality of fan blades and being arranged to be driven byan output from the core shaft.

A fuel tank of the one or more tanks may contain pure SAF having adensity of between 90% and 98% of the density, ρ_(k), of kerosene and acalorific value of between 101% and 105% the calorific value, CV_(K), ofkerosene. Alternatively or additionally, a fuel tank of the one or moretanks may contain a blended fuel comprising a proportion of SAF, the SAFhaving a density of between 90% and 98% of the density, ρ_(k), ofkerosene and a calorific value of between 101% and 105% the calorificvalue, CV_(K), of kerosene and being mixed with a kerosene-based fuel toform the blend.

According to a tenth aspect, there is provided a propulsion system foran aircraft, the propulsion system comprising a gas turbine engine andat least one fuel tank arranged to store fuel to power the gas turbineengine, wherein the stored fuel comprises at least a proportion of asustainable aviation fuel—SAF, the SAF having a density of between 90%and 98% of the density, ρ_(K), of kerosene and a calorific value ofbetween 101% and 105% the calorific value CV_(K), of kerosene,

-   -   the gas turbine engine comprising:        -   a combustor; and        -   a fuel pump arranged to supply fuel from one or more of the            fuel tanks to the combustor, the fuel pump having an inlet            arranged to receive fuel from one or more of the fuel tanks,            the fuel supplied to the pump comprising X % SAF (with any            remainder of the fuel being kerosene) and an outlet arranged            to output the fuel, a proportion of the output fuel being            provided to the combustor and the remainder being            recirculated, the percentage of fuel passing through the            pump which is recirculated/which is not provided to the            combustor being referred to as a spill percentage,            and wherein the propulsion system is arranged such that:    -   the fuel-change spill ratio, R_(s), of:

${R_{s} = \frac{{spill}{percentage}{at}{cruise}{using}{kerosene}}{{spill}{percentage}{at}{cruise}{using}{the}{fuel}{with}X\%{SAF}}},$

-   -   where X % is at least 30%, is greater than or equal to 1.003.

Here, a gravimetric SAF percentage is used, as opposed to volumetric—thefuel is therefore X % SAF by weight. Due to the differing densities ofthe fuels, the volumetric SAF percentage would be slightly different—apercentage by volume may be used in some implementations, and thenumbers give may be adjusted accordingly.

The propulsion system, and more specifically the gas turbine engine, maybe arranged such that:

$R_{s} \geq {1 + {\frac{X}{10000}.}}$

When X is 50 (i.e. the fuel is 50% SAF by weight), the fuel-change spillratio may be at least 1.0066.

When X is 100, such that the fuel is pure SAF, the fuel-change spillratio may be at least 1.0138.

The gas turbine engine may be arranged such that R_(s) varies based onone or more of ambient temperature, altitude, and stage of cruise.

The gas turbine engine may be arranged such that R_(s) decreases by lessthan 0.15% between the beginning and end of cruise, for a constanttemperature and altitude.

The gas turbine engine may be arranged such that R_(s) decreases by atleast 0.11% when altitude increases by at least 600 m.

The gas turbine engine may be arranged such that R_(s) is less than orequal to 1.04. The gas turbine engine may be arranged such that R_(s) isgreater than or equal to 1.003.

The gas turbine engine may be arranged such that R_(s) is greater thanor equal to 1.014.

According to an eleventh aspect, there is provided a method of operatingan aircraft comprising a propulsion system, the propulsion systemcomprising a gas turbine engine including a combustor, and at least onefuel tank arranged to store fuel to power the gas turbine engine,wherein the stored fuel comprises at least a proportion of a sustainableaviation fuel—SAF—having a density of between 90% and 98% of thedensity, ρ_(K), of kerosene and a calorific value of between 101% and105% the calorific value CV_(K), of kerosene, the gas turbine enginecomprising a fuel pump arranged to supply a fuel from one or more of thefuel tanks to the combustor at an energy flow rate, C, the fuel pumphaving an inlet arranged to receive fuel from the one or more fuel tanksand an outlet arranged to output fuel at a pump output volumetric flowrate, Q, a proportion of the output fuel being provided to the combustorand the remainder being recirculated, the percentage of fuel passingthrough the pump which is recirculated being referred to as a spillpercentage, the method comprising:

-   -   supplying a fuel from one or more of the fuel tanks to the gas        turbine engine, the fuel supplied to the gas turbine engine        comprising X % SAF, where X % is in the range from 5% to 100%,        with any remainder of the fuel being kerosene, and wherein the        fuel has a density, ρ_(F), and a calorific value CV_(F); and    -   controlling the propulsion system such that:        -   the fuel-change spill ratio, R_(s), of:

$R_{s} = \frac{{spill}{percentage}{at}{cruise}{using}{kerosene}}{{spill}{percentage}{at}{cruise}{using}{the}{fuel}{with}X\%{SAF}}$

-   -   is equal to:

$\frac{Q - \left( \frac{C}{\left( {{CV}_{K} \times \rho_{K}} \right)} \right)}{Q - \left( \frac{C}{\left( {{CV}_{F} \times \rho_{F}} \right)} \right)}.$

The method may comprise controlling the propulsion system such that anyof the conditions relating to R_(s) as described above for the ninth andtenth aspects apply.

X % may be in the range from 50% to 100%.

According to a twelfth aspect, there is provided a method of operatingan aircraft comprising a propulsion system comprising a gas turbineengine and one or more fuel tanks arranged to contain fuel to power thegas turbine engine, one or more of the tanks containing at least aproportion of a sustainable aviation fuel—SAF—having a density ofbetween 90% and 98% of the density, ρ_(K), of kerosene and a calorificvalue of between 101% and 105% the calorific value CV_(K), of kerosene:

-   -   the gas turbine engine comprising:        -   a combustor; and        -   a fuel pump arranged to supply a fuel from one or more of            the fuel tanks to the combustor, the fuel pump having an            inlet arranged to receive fuel from the one or more fuel            tanks, the fuel supplied to the pump comprising X % SAF,            with any remainder of the fuel being kerosene, and an outlet            arranged to output the fuel, a proportion of the output fuel            being provided to the combustor and the remainder being            recirculated, the percentage of fuel passing through the            pump which is recirculated being referred to as a spill            percentage.

The propulsion system is arranged such that:

-   -   the fuel-change spill ratio, R_(s), is defined as:

${R_{s} = \frac{{spill}{percentage}{at}{cruise}{using}{kerosene}}{{spill}{percentage}{at}{cruise}{using}a{fuel}{with}{}X\%{SAF}}},$

When X % is at least 30% (i.e. when the fuel is at least 30% SAF byweight/mass), R_(s) is greater than or equal to 1.003.

The method may comprise controlling the propulsion system such that anyof the conditions relating to R_(s) as described above for the ninth andtenth aspects apply.

According to a thirteenth aspect, there is provided a method ofdetermining one or more fuel characteristics of an aviation fuelsuitable for powering a gas turbine engine of an aircraft, the methodcomprising:

-   -   determining a mass of a fuel being loaded, or which has been        loaded, onto the aircraft;    -   determining a corresponding volume of the fuel; and    -   determining one or more fuel characteristics of the fuel based        on the determined mass and volume.

The inventors have appreciated that by measuring a mass and volumeparameter of the fuel being loaded, or which has been loaded, onto anaircraft one or more characteristics of the fuel can be determined.

Determining the one or more fuel characteristics may comprisecalculating a fuel density based on the determined fuel mass and fuelvolume. The one or more fuel characteristics may be based on thedensity. The one or more fuel characteristics may be obtained bycomparison of the density to a look-up table of fuels having knowndensities and corresponding fuel characteristics that are to bedetermined.

Determining the mass of the fuel may comprise measuring a mass flow rateat a point within a fuel supply line used to convey fuel to one or morefuel tanks on board the aircraft.

Determining the volume of the fuel may comprise measuring a volume flowrate at a point within the fuel supply line used to convey fuel to oneor more fuel tanks on board the aircraft.

Determining the mass of the fuel may comprise measuring the mass and/ora change in the mass of any one or more of: the aircraft; one or morefuel tanks on board the aircraft; a fuel tanker vehicle from which thefuel is supplied; or a storage vessel from which the fuel is supplied tothe aircraft.

Determining the volume of the fuel may comprise measuring the volumeand/or a change in the volume of fuel: stored in one or more fuel tankson board the aircraft; and/or stored in a fuel storage vessel from whichthe fuel is supplied to the aircraft.

The one or more fuel characteristics determined may include any one ormore of:

-   -   (i) a hydrocarbon distribution of the fuel    -   (ii) a percentage of sustainable aviation fuel in the fuel;        and/or    -   (iii) an aromatic hydrocarbon content of the fuel.

The one or more fuel characteristics may include an indication that thefuel is a fossil fuel e.g. kerosene.

The one or more fuel characteristics may be further determined based ona signal indicative of the temperature of the fuel.

According to a fourteenth aspect, there is provided a fuelcharacteristic determination system for determining one or more fuelcharacteristic of an aviation fuel suitable for powering a gas turbineengine of an aircraft, the system comprising:

-   -   a fuel characteristic determination module arranged to:        -   receive a fuel mass signal indicative of a mass of the fuel            being loaded, or which has been loaded, onto the aircraft;        -   receive a fuel volume signal indicative of a volume of fuel            being loaded or loaded onto the aircraft; and        -   determine one or more fuel characteristics of the fuel based            on the fuel volume and fuel mass signals.

The fuel characteristic determination module may be configured tocalculate a density of the fuel based on the fuel mass signal and fuelvolume signal.

The fuel characteristic determination system may further comprise a masssensor arranged to measure a mass of the fuel, wherein the fuel masssignal is received from the mass sensor.

The fuel characteristic determination system may further comprise avolume sensor arranged to measure a volume of fuel, wherein the fuelvolume signal is received from the volume sensor.

The mass sensor may be a mass flow rate meter. The volume sensor may bea volume flow rate meter.

The mass flow rate meter may be arranged to measure mass flow rate at apoint within a fuel supply line used to convey fuel to one or more fueltanks on board the aircraft.

The volume flow rate meter may be arranged to measure a volume flow rateat a point within the fuel supply line used to supply fuel to one ormore fuel tanks on board the aircraft.

The fuel mass signal may be based on a measured mass and/or change inthe mass of any one or more of: the aircraft; one or more fuel tanks onboard the aircraft; a fuel tanker vehicle from which the fuel issupplied; or a storage vessel from which the fuel is supplied to theaircraft. The mass sensor may be arranged to measure any of theparameter in the previous sentence.

The fuel volume signal may be based on a measured volume and/or inchange in volume of fuel: stored in one or more fuel tanks on board theaircraft; and/or stored in a fuel storage vessel from which the fuel issupplied to the aircraft. The volume sensor may be arranged to measureeither of these parameters.

The one or more fuel characteristics determined may include any one ormore of:

-   -   (i) a hydrocarbon distribution of the fuel;    -   (ii) a percentage of sustainable aviation fuel in the fuel;        and/or    -   (iii) an aromatic hydrocarbon content of the fuel.

The one or more fuel characteristics may include an indication that thefuel is a fossil fuel e.g. kerosene.

The determination module may be further arranged to determine the one ormore fuel characteristics based on a signal indicative of thetemperature of the fuel.

According to a fifteenth aspect, there is provided a method of operatingan aircraft having a gas turbine engine, the method comprising:

-   -   determining one or more fuel characteristics using the method of        the thirteenth aspect; and    -   operating the aircraft according to the one or more fuel        characteristics.

Operating the aircraft according to the one or more fuel characteristicsmay comprise:

-   -   a) modifying a control parameter of the aircraft, preferably a        control parameter of the gas turbine engine, in response to the        one or more fuel characteristics; and/or    -   b) providing a fuel having different fuel characteristics during        refuelling of the aircraft.

According to a sixteenth aspect, there is provided an aircraftcomprising the fuel characteristic determination system of thefourteenth aspect, the aircraft further comprising a control systemarranged to control operation of the aircraft according to the one ormore fuel characteristics determined by the fuel characteristicdetermination system.

Accord to a seventeenth aspect, there is provided a method ofdetermining one or more fuel characteristics of an aviation fuelsuitable for powering a gas turbine engine of an aircraft, the gasturbine engine having a combustor supplied with fuel from a fuel systemof the aircraft, the method comprising:

-   -   determining a mass of the fuel being supplied to the combustor;    -   determining a corresponding volume of the fuel being supplied to        the combustor; and    -   determining one or more fuel characteristics based on the        determined mass and volume.

The inventors have appreciated that one or more fuel characteristics canbe determined during operation of a gas turbine engine by measuring amass and volume parameter of fuel as it is being supplied to a combustorof the gas turbine engine. One or more fuel characteristics can bedetermined based on the measure mass and volume of the fuel.

Determining the one or more fuel characteristics may comprisecalculating a fuel density based on the determined fuel mass and fuelvolume. The one or more fuel characteristics may be based on thedensity. The one or more fuel characteristics may be obtained bycomparison of the density to a look-up table of fuels having knowndensities and corresponding fuel characteristics that are to bedetermined.

Determining the mass of the fuel may comprise determining a mass flowrate of fuel being supplied to the combustor.

The fuel system may comprise an engine fuel system forming part of thegas turbine engine. The mass flow rate may be measured at a point in afuel conduit of the engine fuel system. The mass flow rate may bemeasured immediately before fuel reaches the combustor.

The mass flow rate may be determined based on an operating parameter ofa fuel pump provided in the fuel system. The mass flow rate may bedetermined based on a measurement of fuel flow using a mass flow meter.

Determining the volume of the fuel may comprise determining a volumeflow rate of fuel being supplied to the combustor.

The volume flow rate may be measured at a point in a fuel conduit of theengine fuel system. The volume flow rate may be measured immediatelybefore fuel reaches the combustor. The volume flow rate may be measuredat a position adjacent the mass flow rate.

The volume flow rate may be determined based on an operating parameterof a pump provided in the fuel system. The volume flow rate may bedetermined based on a measurement of fuel flow using a volume flowmeter.

The one or more fuel characteristics determined include any one or moreof:

-   -   (i) a hydrocarbon distribution of the fuel;    -   (ii) a percentage of sustainable aviation fuel in the fuel;        and/or    -   (iii) an aromatic hydrocarbon content of the fuel.

The one or more fuel characteristics include an indication that the fuelis a fossil fuel e.g. kerosene.

According to an eighteenth aspect, there is provided a fuelcharacteristic determination system for determining one or more fuelcharacteristics of an aviation fuel suitable for powering a gas turbineengine of an aircraft, the gas turbine engine having a combustorsupplied with fuel from a fuel system of the aircraft, the fuelcharacteristic determination system comprising:

-   -   a fuel characteristic determination module arranged to:        -   receive a fuel mass signal indicative of a mass of fuel            being supplied to the combustor;        -   receive a fuel volume signal indicative of a volume of fuel            being supplied to the combustor; and        -   determine one or more fuel characteristics of the fuel based            on the fuel volume and fuel mass signals.

The fuel characteristic determination module may be configured tocalculate a density of the fuel based on the fuel mass signal and fuelvolume signal.

The fuel characteristic determination system may further comprise a massflow meter arranged to measure a mass flow rate of fuel being suppliedto the combustor. The fuel mass signal may be received from the massflow meter.

The fuel characteristic determination system may further comprise avolume flow meter arranged to measure a volume flow rate of fuel beingsupplied to the combustor. The fuel volume signal may be received fromthe volume flow meter.

The mass flow meter and/or volume flow meter may be arranged to measurethe flow of fuel measured at a point in a fuel conduit of an engine fuelsystem of the gas turbine engine. The mass flow meter and/or volume flowmeter may be arranged to measure the flow of fuel measured at a pointimmediately before fuel reaches the combustor.

The fuel system may further comprise a fuel pump arranged to providefuel to the combustor. The fuel mass signal and/or the fuel volumesignal may be based on an operating parameter of the fuel pump.

The one or more fuel characteristics determined include any one or moreof:

-   -   (i) a hydrocarbon distribution of the fuel;    -   (ii) a percentage of sustainable aviation fuel in the fuel;        and/or    -   (iii) an aromatic hydrocarbon content of the fuel.

The one or more fuel characteristics include an indication that the fuelis a fossil fuel e.g. kerosene.

According to a nineteenth aspect, there is provided a method ofoperating an aircraft having a gas turbine engine, the methodcomprising:

-   -   determining one or more fuel characteristics using the method of        the seventeenth aspect; and    -   operating the aircraft according to the one or more fuel        characteristics.

Operating the aircraft according to the one or more fuel characteristicsmay comprise:

-   -   a) modifying a control parameter of the aircraft, preferably a        control parameter of the gas turbine engine, in response to the        one or more fuel characteristics; and/or    -   b) providing a fuel having different fuel characteristics during        refuelling of the aircraft.

According to a twentieth aspect, there is provided an aircraftcomprising the fuel characteristic determination system of theeighteenth aspect, the aircraft further comprising a control systemarranged to control operation of the aircraft according to the one ormore fuel characteristics determined by the fuel characteristicdetermination system.

The present disclosure may apply to any relevant configuration of gasturbine engine. Such a gas turbine engine may be, for example, aturbofan gas turbine engine, an open rotor gas turbine engine (in whichthe propeller is not surrounded by a nacelle), a turboprop engine or aturbojet engine. Any such engine may or may not be provided with anafterburner.

A gas turbine engine in accordance with any aspect of the presentdisclosure may comprise an engine core comprising a turbine, acombustor, a compressor, and a core shaft connecting the turbine to thecompressor. Such a gas turbine engine may comprise a fan (having fanblades). Such a fan may be located upstream of the engine core.Alternatively, in some examples, the gas turbine engine may comprise afan located downstream of the engine core, for example where the gasturbine engine is an open rotor or a turboprop engine (in which case thefan may be referred to as a propeller).

Where the gas turbine engine is an open rotor or a turboprop engine, thegas turbine engine may comprise two contra-rotating propeller stagesattached to and driven by a free power turbine via a shaft. Thepropellers may rotate in opposite senses so that one rotates clockwiseand the other anti-clockwise around the engine's rotational axis.Alternatively, the gas turbine engine may comprise a propeller stage anda guide vane stage configured downstream of the propeller stage. Theguide vane stage may be of variable pitch. Accordingly, high-pressure,intermediate pressure, and free power turbines respectively may drivehigh and intermediate pressure compressors and propellers by suitableinterconnecting shafts. Thus, the propellers may provide the majority ofthe propulsive thrust.

Where the gas turbine engine is an open rotor or a turboprop engine, oneor more of the propellor stages may be driven by a gearbox. The gearboxmay be of the type described herein.

An engine according to the present disclosure may be a turbofan engine.Such an engine may be a direct-drive turbofan engine in which the fan isdirectly connected to the fan drive turbine, for example without agearbox. In such a direct-drive turbofan engine, the fan may be said torotate at the same rotational speed as the fan-drive turbine.

An engine according to the present disclosure may be a geared turbofanengine. In such an arrangement, the engine has a fan that is driven viaa gearbox. Accordingly, such a gas turbine engine may comprise a gearboxthat receives an input from the core shaft and outputs drive to the fanso as to drive the fan at a lower rotational speed than the core shaft.The input to the gearbox may be directly from the core shaft, orindirectly from the core shaft, for example via a spur shaft and/orgear. The core shaft may rigidly connect the turbine and the compressor,such that the turbine and compressor rotate at the same speed (with thefan rotating at a lower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan isa lower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Such agearbox may be a single stage. Alternatively, such a gearbox may be acompound gearbox, for example a compound planetary gearbox (which mayhave the input on the sun gear and the output on the ring gear, and thusbe referred to as a “compound star” gearbox), for example having twostages of reduction.

The gearbox may have any desired reduction ratio (defined as therotational speed of the input shaft divided by the rotational speed ofthe output shaft), for example greater than 2.5, for example in therange of from 3 to 4.2, or 3.2 to 3.8, for example on the order of or atleast 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. Thegear ratio may be, for example, between any two of the values in theprevious sentence. Purely by way of example, the gearbox may be a “star”gearbox having a reduction ratio in the range of from 3.1 or 3.2 to 3.8.Purely by way of further example, the gearbox may be a “star” gearboxhaving a reduction ratio in the range 3.0 to 3.1. Purely by way offurther example, the gearbox may be a “planetary” gearbox having areduction ratio in the range 3.6 to 4.2. In some arrangements, the gearratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, fuel of agiven composition or blend is provided to a combustor, which may beprovided downstream of the fan and compressor(s) with respect to theflowpath (for example axially downstream). For example, the combustormay be directly downstream of (for example at the exit of) the secondcompressor, where a second compressor is provided. By way of furtherexample, the flow at the exit to the combustor may be provided to theinlet of the second turbine, where a second turbine is provided. Thecombustor may be provided upstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other. Forexample, the gas turbine engine may be a direct-drive turbofan gasturbine engine comprising 13 or 14 compressor stages (in addition to thefan). Such an engine may, for example, comprise 3 stages in the first(or “low pressure”) compressor and either 10 or 11 stages in the second(or “high pressure”) compressor. By way of further example, the gasturbine engine may be a “geared” gas turbine engine (in which the fan isdrive by a first core shaft via a reduction gearbox) comprising 11, 12or 13 compressor stages (in addition to the fan). Such an engine maycomprise 3 or 4 stages in the first (or “low pressure”) compressor and 8or 9 stages in the second (or “high pressure”) compressor. By way offurther example, the gas turbine engine may be a “geared” gas turbineengine having 4 stages in the first (or “low pressure”) compressor and10 stages in the second (or “high pressure”) compressor.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other. The second (or “high pressure”) turbinemay comprise 2 stages in any arrangement (for example regardless ofwhether it is a geared or direct-drive engine). The gas turbine enginemay be a direct-drive gas turbine engine comprising a first (or “lowpressure”) turbine having 5, 6 or 7 stages. Alternatively, the gasturbine engine may be a “geared” gas turbine engine comprising a first(or “low pressure”) turbine having 3 or 4 stages.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38, 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32 or 0.29 to 0.30.These ratios may commonly be referred to as the hub-to-tip ratio. Theradius at the hub and the radius at the tip may both be measured at theleading edge (or axially forwardmost) part of the blade. The hub-to-tipratio refers, of course, to the gas-washed portion of the fan blade,i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 140 cm, 170 cm, 180 cm, 190 cm, 200 cm, 210 cm, 220cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm(around 130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm, 390cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm(around 165 inches). The fan diameter may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds), for example in the range of from240 cm to 280 cm or 330 cm to 380 cm. Purely by way of non-limitativeexample, the fan diameter may be in the range of from 170 cm to 180 cm,190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 230 cm, 290 cm to 300 cmor 340 cm to 360 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 3500 rpm, for example less than 2500 rpm,for example less than 2300 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for an“geared” gas turbine engine having a fan diameter in the range of from200 cm to 210 cm may be in the range of from 2750 to 2900 rpm. Purely byway of further non-limitative example, the rotational speed of the fanat cruise conditions for an “geared” gas turbine engine having a fandiameter in the range of from 210 cm to 230 cm may be in the range offrom 2500 to 2800 rpm. Purely by way of further non-limitative example,the rotational speed of the fan at cruise conditions for an “geared” gasturbine engine having a fan diameter in the range of from 340 cm to 360cm may be in the range of from 1500 to 1800 rpm. . Purely by way offurther non-limitative example, the rotational speed of the fan atcruise conditions for a direct drive engine having a fan diameter in therange of from 190 cm to 200 cm may be in the range of from 3600 to 3900rpm. Purely by way of further non-limitative example, the rotationalspeed of the fan at cruise conditions for a direct drive engine having afan diameter in the range of from 300 cm to 340 cm may be in the rangeof from 2000 to 2800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 23 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (allvalues being dimensionless). The fan tip loading may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds), for example in the range offrom 0.28 to 0.31, or 0.29 to 0.3 (for example for a geared gas turbineengine).

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 9.

9.5, 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16,16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratio may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 12 to 16, 13 to 15, or 13 to 14. Purely by way ofnon-limitative example, the bypass ratio of a direct-drive gas turbineengine according to the present disclosure may be in the range of from9:1 to 11:1. Purely by way of further non-limitative example, the bypassratio of a geared gas turbine engine according to the present disclosuremay be in the range of from 12:1 to 15:1 The bypass duct may besubstantially annular. The bypass duct may be radially outside the coreengine. The radially outer surface of the bypass duct may be defined bya nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressure atthe exit of the highest pressure compressor (before entry into thecombustor) to the stagnation pressure upstream of the fan. By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 50 to 70.Purely by way of non-limitative example, the overall pressure ratio atcruise conditions of a geared gas turbine engine having a fan diameterin the range of from 200 cm to 210 cm may be in the range of from 40 to45. Purely by way of non-limitative example, the overall pressure ratioat cruise conditions of a geared gas turbine engine having a fandiameter in the range of from 210 cm to 230 cm may be in the range offrom 45 to 55. Purely by way of non-limitative example, the overallpressure ratio at cruise conditions of a geared gas turbine enginehaving a fan diameter in the range of from 340 cm to 360 cm may be inthe range of from 50 to 60. Purely by way of non-limitative example, theoverall pressure ratio at cruise conditions of a direct-drive gasturbine engine having a fan diameter in the range of from 300 cm to 340cm may be in the range of from 50 to 60.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. In someexamples, specific thrust may depend, for a given thrust condition, uponthe specific composition of fuel provided to the combustor. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹ s, 105 Nkg⁻¹ s, 100 Nkg⁻¹ s, 95 Nkg⁻¹ s, 90 Nkg⁻¹ s, 85 Nkg⁻¹ s,or 80 Nkg⁻¹ s. The specific thrust may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 80 Nkg⁻¹ sto 100 Nkg⁻¹ s, or 85 Nkg⁻¹ s to 95 Nkg⁻¹ s. Such engines may beparticularly efficient in comparison with conventional gas turbineengines. Purely by way of non-limitative example, the specific thrust ofa geared gas turbine engine having a fan diameter in the range of from200 cm to 210 cm may be in the range of from 90 Nkg⁻¹ s to 95 Nkg⁻¹ s.Purely by way of non-limitative example, the specific thrust of a gearedgas turbine engine having a fan diameter in the range of from 210 cm to230 cm may be in the range of from 80 Nkg⁻¹ s to 90 Nkg⁻¹ s. Purely byway of non-limitative example, the specific thrust of a geared gasturbine engine having a fan diameter in the range of from 340 cm to 360cm may be in the range of from 70 Nkg⁻¹ s to 90 Nkg⁻¹ s. Purely by wayof non-limitative example, the specific thrust of a direct drive gasturbine engine having a fan diameter in the range of from 300 cm to 340cm may be in the range of from 90 Nkg⁻¹ s to 120 Nkg⁻¹ s.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:100 kN, 110 kN, 120 kN, 130 kN, 140 kN, 150 kN, 160 kN, 170 kN, 180 kN,190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550kN. The maximum thrust may be in an inclusive range bounded by any twoof the values in the previous sentence (i.e. the values may form upperor lower bounds). Purely by way of non-limitative example, a gas turbineas described and/or claimed herein may be capable of producing a maximumthrust in the range of from 330 kN to 420 kN, for example 350kN to 400kN. Purely by way of non-limitative example, the maximum thrust of ageared gas turbine engine having a fan diameter in the range of from 200cm to 210 cm may be in the range of from 140 kN to 160 kN. Purely by wayof non-limitative example, the maximum thrust of a geared gas turbineengine having a fan diameter in the range of from 210 cm to 230 cm maybe in the range of from 150 kN to 200 kN. Purely by way ofnon-limitative example, the maximum thrust of a geared gas turbineengine having a fan diameter in the range of from 340 cm to 360 cm maybe in the range of from 370 kN to 500 kN. Purely by way ofnon-limitative example, the maximum thrust of a direct drive gas turbineengine having a fan diameter in the range of from 300 cm to 340 cm maybe in the range of from 370 kN to 500 kN. The thrust referred to abovemay be the maximum net thrust at standard atmospheric conditions at sealevel plus 15 degrees C. (ambient pressure 101.3 kPa, temperature 30degrees C.), with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. In some examples, TET may depend,for a given thrust condition, upon the specific composition of fuelprovided to the combustor. At cruise, the TET may be at least (or on theorder of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or1650K. Thus, purely by way of non-limitative example, the TET at cruiseof a geared gas turbine engine having a fan diameter in the range offrom 200 cm to 210 cm may be in the range of from 1540K to 1600K. Purelyby way of non-limitative example, the TET at cruise of a geared gasturbine engine having a fan diameter in the range of from 210 cm to 230cm may be in the range of from 1590K to 1650K. Purely by way ofnon-limitative example, the TET at cruise of a geared gas turbine enginehaving a fan diameter in the range of from 340 cm to 360 cm may be inthe range of from 1600K to 1660K. Purely by way of non-limitativeexample, the TET at cruise of a direct drive gas turbine engine having afan diameter in the range of from 300 cm to 340 cm may be in the rangeof from 1590K to 1650K. Purely by way of non-limitative example, the TETat cruise of a direct drive gas turbine engine having a fan diameter inthe range of from 300 cm to 340 cm may be in the range of from 1570K to1630K.

The TET at cruise may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). The maximum TET in use of the engine may be, for example, atleast (or on the order of) any of the following: 1700K, 1750K, 1800K,1850K, 1900K, 1950K, 2000K, 2050K, or 2100K. Thus, purely by way ofnon-limitative example, the maximum TET of a geared gas turbine enginehaving a fan diameter in the range of from 200 cm to 210 cm may be inthe range of from 1890K to 1960K. Purely by way of non-limitativeexample, the maximum TET of a geared gas turbine engine having a fandiameter in the range of from 210 cm to 230 cm may be in the range offrom 1890K to 1960K. Purely by way of non-limitative example, themaximum TET of a geared gas turbine engine having a fan diameter in therange of from 340 cm to 360 cm may be in the range of from 1890K to1960K. Purely by way of non-limitative example, the maximum TET of adirect drive gas turbine engine having a fan diameter in the range offrom 300 cm to 340 cm may be in the range of from 1935K to 1995K. Purelyby way of non-limitative example, the maximum TET of a direct drive gasturbine engine having a fan diameter in the range of from 300 cm to 340cm may be in the range of from 1890K to 1950K. The maximum TET may be inan inclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 1800K to 1950K. The maximum TET may occur, forexample, at a high thrust condition, for example at a maximum take-off(MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre composite. By way of further example at least apart of the fan blade and/or aerofoil may be manufactured at least inpart from a metal, such as a titanium based metal or an aluminium basedmaterial (such as an aluminium-lithium alloy) or a steel based material.The fan blade may comprise at least two regions manufactured usingdifferent materials. For example, the fan blade may have a protectiveleading edge, which may be manufactured using a material that is betterable to resist impact (for example from birds, ice or other material)than the rest of the blade. Such a leading edge may, for example, bemanufactured using titanium or a titanium-based alloy. Thus, purely byway of example, the fan blade may have a carbon-fibre or aluminium basedbody (such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmay be formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades. Where the fan blades have a carbon fibre composite body,there may be 16 or 18 fan blades. Where the fan blades have a metallicbody (for example aluminium-lithium or titanium-alloy), there may be 18,20 or 22 fan blades.

As used herein, the terms idle, taxi, take-off, climb, cruise, descent,approach, and landing have the conventional meaning and would be readilyunderstood by the skilled person. Thus, for a given gas turbine enginefor an aircraft, the skilled person would immediately recognise eachterm to refer to an operating phase of the engine within a given missionof an aircraft to which the gas turbine engine is designed to beattached.

In this regard, ground idle may refer to an operating phase of theengine where the aircraft is stationary and in contact with the ground,but where there is a requirement for the engine to be running. Duringidle, the engine may be producing between 3% and 9% of the availablethrust of the engine. In further non-limitative examples, the engine maybe producing between 5% and 8% of available thrust. In furthernon-limitative examples, the engine may be producing between 6% and 7%of available thrust. Taxi may refer to an operating phase of the enginewhere the aircraft is being propelled along the ground by the thrustproduced by the engine. During taxi, the engine may be producing between5% and 15% of available thrust. In further non-limitative examples, theengine may be producing between 6% and 12% of available thrust. Infurther non-limitative examples, the engine may be producing between 7%and 10% of available thrust. Take-off may refer to an operating phase ofthe engine where the aircraft is being propelled by the thrust producedby the engine. At an initial stage within the take-off phase, theaircraft may be propelled whilst the aircraft is in contact with theground. At a later stage within the take-off phase, the aircraft may bepropelled whilst the aircraft is not in contact with the ground. Duringtake-off, the engine may be producing between 90% and 100% of availablethrust. In further non-limitative examples, the engine may be producingbetween 95% and 100% of available thrust. In further non-limitativeexamples, the engine may be producing 100% of available thrust.

Climb may refer to an operating phase of the engine where the aircraftis being propelled by the thrust produced by the engine. During climb,the engine may be producing between 75% and 100% of available thrust. Infurther non-limitative examples, the engine may be producing between 80%and 95% of available thrust. In further non-limitative examples, theengine may be producing between 85% and 90% of available thrust. In thisregard, climb may refer to an operating phase within an aircraft flightcycle between take-off and the arrival at cruise conditions.Additionally or alternatively, climb may refer to a nominal point in anaircraft flight cycle between take-off and landing, where a relativeincrease in altitude is required, which may require an additional thrustdemand of the engine.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/or distance -between top of climb and start of descent. Cruise conditions thus definean operating point of the gas turbine engine that provides a thrust thatwould ensure steady state operation (i.e. maintaining a constantaltitude and constant Mach Number) at mid-cruise of an aircraft to whichit is designed to be attached, taking into account the number of enginesprovided to that aircraft. For example where an engine is designed to beattached to an aircraft that has two engines of the same type, at cruiseconditions the engine provides half of the total thrust that would berequired for steady state operation of that aircraft at mid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000m, for example in the range of from 10000 m to 12000 m, for examplein the range of from 10400 m to 11600 m (around 38000 ft), for examplein the range of from 10500 m to 11500 m, for example in the range offrom 10600 m to 11400 m, for example in the range of from 10700 m(around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

Furthermore, the skilled person would immediately recognise either orboth of descent and approach to refer to an operating phase within anaircraft flight cycle between cruise and landing of the aircraft. Duringeither or both of descent and approach, the engine may be producingbetween 20% and 50% of available thrust. In further non-limitativeexamples, the engine may be producing between 25% and 40% of availablethrust. In further non-limitative examples, the engine may be producingbetween 30% and 35% of available thrust. Additionally or alternatively,descent may refer to a nominal point in an aircraft flight cycle betweentake-off and landing, where a relative decrease in altitude is required,and which may require a reduced thrust demand of the engine.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat any suitable condition, which may be as defined elsewhere herein (forexample in terms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at any suitable condition, for example the mid-cruise of theaircraft, as defined elsewhere herein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is a schematic view of an aircraft including a refuellingmanager;

FIG. 5 is a schematic representation of a refuelling management method;

FIG. 6 is a schematic view of an aircraft including a fuel trackingsystem;

FIG. 7 is a schematic representation of a calorific value checkingmethod;

FIG. 8 is a schematic representation of a fuel tracking system;

FIG. 9 is a schematic representation of a refuelling management methodincluding the calorific value checking method of FIG. 7 ;

FIG. 10 is a schematic view of an aircraft including a fuel trackingsystem;

FIG. 11 is a schematic representation of a calorific value determiningmethod;

FIG. 12 is a schematic representation of a fuel tracking system;

FIG. 13 is a schematic view of an aircraft fuel supply system,illustrating spill around the fuel pump;

FIG. 14 is a schematic view of an aircraft fuel pump;

FIG. 15 is a schematic representation of a spill management method; and

FIG. 16 is a schematic view of an aircraft including a fuelcharacteristic determination system;

FIG. 17 is a schematic representation of a method of determining one ormore fuel characteristics of an aviation fuel;

FIG. 18 is schematic view of a fuel characteristic determination systemprovided within a fuel system of a gas turbine engine;

FIG. 19 is another schematic representation of a method of determiningone or more fuel characteristics of an aviation fuel; and

FIG. 20 is a schematic representation of a method of operating anaircraft.

DETAILED DESCRIPTION OF THE DISCLOSURE

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel F and the mixture is combusted.The combustion equipment 16 may be referred to as the combustor 16, withthe terms “combustion equipment 16” and “combustor 16” usedinterchangeably herein. The resultant hot combustion products thenexpand through, and thereby drive, the high pressure and low pressureturbines 17, 19 before being exhausted through the nozzle 20 to providesome propulsive thrust. The high pressure turbine 17 drives the highpressure compressor 15 by a suitable interconnecting shaft 27. The fan23 generally provides the majority of the propulsive thrust. Theepicyclic gearbox 30 is a reduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2 . The low pressure turbine 19 (see FIG. 1 ) drives the shaft26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclicgear arrangement 30. Radially outwardly of the sun gear 28 andintermeshing therewith is a plurality of planet gears 32 that arecoupled together by a planet carrier 34. The planet carrier 34constrains the planet gears 32 to precess around the sun gear 28 insynchronicity whilst enabling each planet gear 32 to rotate about itsown axis. The planet carrier 34 is coupled via linkages 36 to the fan 23in order to drive its rotation about the engine axis 9. Radiallyoutwardly of the planet gears 32 and intermeshing therewith is anannulus or ring gear 38 that is coupled, via linkages 40, to astationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e.

not including the fan 23 ) respectively and/or the turbine andcompressor stages that are connected together by the interconnectingshaft 26 with the lowest rotational speed in the engine (i.e. notincluding the gearbox output shaft that drives the fan 23). In someliterature, the “low pressure turbine” and “low pressure compressor”referred to herein may alternatively be known as the “intermediatepressure turbine” and “intermediate pressure compressor”. Where suchalternative nomenclature is used, the fan 23 may be referred to as afirst, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3 . Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3 . There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2 . For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2 .

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area.

Whilst the described example relates to a turbofan engine, thedisclosure may apply, for example, to any type of gas turbine engine,such as an open rotor (in which the fan stage is not surrounded by anacelle) or turboprop engine, for example. In some arrangements, the gasturbine engine 10 may not comprise a gearbox 30.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1 ), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

The fuel F provided to the combustion equipment 16 may comprise afossil-based hydrocarbon fuel, such as Kerosene. Thus, the fuel F maycomprise molecules from one or more of the chemical families ofn-alkanes, iso-alkanes, cycloalkanes, and aromatics. Additionally oralternatively, the fuel F may comprise renewable hydrocarbons producedfrom biological or non-biological resources, otherwise known assustainable aviation fuel (SAF). In each of the provided examples, thefuel F may comprise one or more trace elements including, for example,sulphur, nitrogen, oxygen, inorganics, and metals.

Functional performance of a given composition, or blend of fuel for usein a given mission, may be defined, at least in part, by the ability ofthe fuel to service the Brayton cycle of the gas turbine engine 10.Parameters defining functional performance may include, for example,specific energy; energy density; thermal stability; and, emissionsincluding particulate matter. A relatively higher specific energy (i.e.energy per unit mass), expressed as MJ/kg, may at least partially reducetake-off weight, thus potentially providing a relative improvement infuel efficiency. A relatively higher energy density (i.e. energy perunit volume), expressed as MJ/L, may at least partially reduce take-offfuel volume, which may be particularly important for volume-limitedmissions or military operations involving refuelling. A relativelyhigher thermal stability (i.e. inhibition of fuel to degrade or cokeunder thermal stress) may permit the fuel to sustain elevatedtemperatures in the engine and fuel injectors, thus potentiallyproviding relative improvements in combustion efficiency. Reducedemissions, including particulate matter, may permit reduced contrailformation, whilst reducing the environmental impact of a given mission.Other properties of the fuel may also be key to functional performance.For example, a relatively lower freeze point (° C.) may allow long-rangemissions to optimise flight profiles; minimum aromatic concentrations(%) may ensure sufficient swelling of certain materials used in theconstruction of o-rings and seals that have been previously exposed tofuels with high aromatic contents; and, a maximum surface tension (mN/m)may ensure sufficient spray break-up and atomisation of the fuel.

The ratio of the number of hydrogen atoms to the number of carbon atomsin a molecule may influence the specific energy of a given composition,or blend of fuel. Fuels with higher ratios of hydrogen atoms to carbonatoms may have higher specific energies in the absence of bond strain.For example, fossil-based hydrocarbon fuels may comprise molecules withapproximately 7 to 18 carbons, with a significant portion of a givencomposition stemming from molecules with 9 to 15 carbons, with anaverage of 12 carbons.

A number of sustainable aviation fuel blends have been approved for use,comprising between 10% and 50% sustainable aviation fuel (the remaindercomprising one or more fossil-based hydrocarbon fuels, such asKerosene), with further compositions awaiting approval. However, thereis an anticipation in the aviation industry that sustainable aviationfuel blends comprising up to (and including) 100% sustainable aviationfuel (SAF) will be eventually approved for use.

Sustainable aviation fuels may comprise one or more of n-alkanes,iso-alkanes, cyclo-alkanes, and aromatics, and may be produced, forexample, from one or more of synthesis gas (syngas); lipids (e.g. fats,oils, and greases); sugars; and alcohols. Thus, sustainable aviationfuels may comprise either or both of a lower aromatic and sulphurcontent, relative to fossil-based hydrocarbon fuels. Additionally oralternatively, sustainable aviation fuels may comprise either or both ofa higher iso-alkane and cyclo-alkane content, relative to fossil-basedhydrocarbon fuels. Thus, in some examples, sustainable aviation fuelsmay comprise either or both of a density of between 90% and 98% that ofkerosene and a calorific value of between 101% and 105% that ofkerosene.

Owing at least in part to the molecular structure of sustainableaviation fuels, sustainable aviation fuels may provide benefitsincluding, for example, one or more of a higher specific energy(despite, in some examples, a lower energy density); higher specificheat capacity; higher thermal stability; higher lubricity; lowerviscosity; lower surface tension; lower freeze point; lower sootemissions; and, lower CO₂ emissions, relative to fossil-basedhydrocarbon fuels (e.g. when combusted in the combustion equipment 16).Accordingly, relative to fossil-based hydrocarbon fuels, such asKerosene, sustainable aviation fuels may lead to either or both of arelative decrease in specific fuel consumption, and a relative decreasein maintenance costs.

As depicted in FIGS. 4, 6 and 10 , an aircraft 1 may comprise multiplefuel tanks 50, 53; for example a larger, primary fuel tank 50 located inthe aircraft fuselage, and a smaller fuel tank 53 a, 53 b located ineach wing. In other examples, an aircraft 1 may have only a single fueltank 50, and/or the wing fuel tanks 53 may be larger than the centralfuel tank 50, or no central fuel tank may be provided (with all fuelinstead being stored in the aircraft's wings)—it will be appreciatedthat many different tank layouts are envisaged and that the examplespictured are provided for ease of description and not intended to belimiting.

FIG. 4 , FIG. 6 , and FIG. 10 show an aircraft 1 with a propulsionsystem 2 comprising two gas turbine engines 10. The gas turbine engines10 are supplied with fuel from a fuel supply system on board theaircraft 1. The fuel supply system of the examples pictured comprises asingle fuel source. For the purposes of the present application the term“fuel source” is understood to mean either 1) a single fuel tank, or 2)a plurality of fuel tanks which are fluidly interconnected. Each of thefuel sources is arranged to provide a separate source of fuel i.e. thefirst fuel source may contain a first fuel having a different fuelcharacteristic, or multiple different fuel characteristics, from asecond fuel contained in the second fuel source. First and second fuelsources are therefore not fluidly coupled to each other so as toseparate the different fuels (at least under normal running conditions).

As used herein, the term “fuel characteristics” refers to intrinsic orinherent fuel properties such as fuel composition, not variableproperties such as volume or temperature. Examples of fuelcharacteristics include one or more of:

-   -   i. the percentage of sustainable aviation fuel (SAF) in the        fuel, or an indication that the fuel is a fossil fuel, for        example fossil kerosene, or a pure SAF;    -   ii. parameters of a hydrocarbon distribution of the fuel, such        as:        -   the aromatic hydrocarbon content of the fuel, and optionally            also / alternatively the multi-aromatic hydrocarbon content            of the fuel;        -   the hydrogen to carbon ratio (H/C) of the fuel; 10%            composition information for some or all hydrocarbons            present;    -   iii. the presence or percentage of a particular element or        species, such as:        -   the percentage of nitrogen-containing species in the fuel;        -   the presence or percentage of a tracer species or trace            element in the fuel (e.g. a trace substance inherently            present in the fuel which may vary between fuels and so be            used to identify a fuel, and/or a substance added            deliberately to act as a tracer);        -   naphthalene content of the fuel;        -   sulphur content of the fuel;        -   cycloparaffin content of the fuel;        -   oxygen content of the fuel;    -   iv. one or more properties of the fuel in use in a gas turbine        engine 10, such as:        -   level of non-volatile particulate matter (nvPM) emissions or            CO₂ emissions on combustion (a value may be provided for a            specific combustor operating under particular conditions to            compare fuels fairly—a measured value may be adjusted            accordingly based on combustor properties and conditions);        -   level of coking of the fuel;    -   v. one or more properties of the fuel itself, independent of use        in an engine or combustion, such as:        -   thermal stability of the fuel (e.g. thermal breakdown            temperature); and        -   one or more physical properties such as density, viscosity,            calorific value, freeze temperature, and/or heat capacity.

In the present example, the first fuel source comprises a centre fueltank 50, located primarily in the fuselage of the aircraft 1 and aplurality of wing fuel tanks 53 a, 53 b, where at least one wing fueltank is located in the port wing and at least one wing fuel tank islocated in the starboard wing for balancing. All of the tanks 50, 53 arefluidly interconnected in the example shown, so forming a single fuelsource. Each of the centre fuel tank 50 and the wing fuel tanks 53 maycomprise a plurality of fluidly interconnected fuel tanks.

In another example, the wing fuel tanks 53 a, 53 b may not be fluidlyconnected to the central tank 50, so forming a separate, second fuelsource. For balancing purposes, one or more fuel tanks in the port wingmay be fluidly connected to one or more fuel tanks in the starboardwing. This may be done either via a/the centre fuel tank 50 (if thattank does not form part of a different fuel source), or bypassing thecentre fuel tank(s), or both (for maximum flexibility and safety).

In another example, the first fuel source comprises wing fuel tanks 53and a centre fuel tank 50, while a second fuel source comprises afurther separate centre fuel tank. Fluid interconnection between wingfuel tanks 53 and the centre fuel tank 50 of the first fuel source maybe provided for balancing of the aircraft 1.

In some examples, the allocation of fuel tanks 50, 53 available on theaircraft 1 may be constrained such that a first fuel source and a secondfuel source are each substantially symmetrical with respect to theaircraft centre line. In cases where an asymmetric fuel tank allocationis permitted, a suitable means of fuel transfer may be provided betweenfuel tanks of the first fuel source and/or between fuel tanks of thesecond fuel source such that the position of the aircraft's centre ofmass can be maintained within acceptable lateral limits throughout theflight.

An aircraft 1 may be refuelled by connecting a fuel storage vessel 60,such as that provided by an airport fuel truck, or a permanent pipeline,to a fuel line connection port 62 of the aircraft, via a fuel line 61. Adesired amount of fuel may be transferred from the fuel storage vessel60 to the one or more tanks 50, 53 of the aircraft 1. Especially inexamples with more than one fuel source, in which different tanks 50, 53are to be filled with different fuels, multiple fuel line connectionports 62 may be provided instead of one, and/or valves may be used todirect fuel appropriately.

Aircraft 1 generally refuel at multiple different airports, for exampleat the beginning and end of a long-distance flight. Whilst there arestandards with which all aviation fuels must be compliant, differentaviation fuels have different compositions, for example depending ontheir source (e.g. different petroleum sources, biofuels or othersynthetic aviation fuels (often described as sustainable aviationfuels—SAFs), and/or mixtures of petroleum-based fuels, and other fuels)and on any additives included (e.g. such as antioxidants and metaldeactivators, biocides, static reducers, icing inhibitors, corrosioninhibitors) and any impurities. As well as varying between airports andfuel suppliers, even for a given airport or fuel supplier, fuelcomposition of the available aviation fuel may vary between batches.Further, fuel tanks 50, 53 of aircraft 1 are usually not emptied beforebeing topped up for a subsequent flight, resulting in mixtures ofdifferent fuels within the tanks—effectively a fuel with a differentcomposition resulting from the mixture. Different fuels may havedifferent calorific values and/or different densities, among other fuelcharacteristics.

As used herein, the term “calorific value” denotes the lower heatingvalue (also known as net calorific value) of the fuel, unless otherwisespecified. The net calorific value is defined as the amount of heatreleased by combusting a specified quantity of the fuel, assuming thatthe latent heat of vaporisation of water in the reaction products is notrecovered (i.e. that produced water remains as water vapour aftercombustion).

Calorific values (also referred to as heating values) of fuels may bedirectly determined - for example by measuring the energy released whena certain volume or mass of the fuel is combusted in the gas turbineengine 10—or calculated from other fuel parameters; e.g. looking at thehydrocarbon distribution of the fuel and the calorific value of eachconstituent hydrocarbon type. Alternatively, or additionally so as toprovide verification, the calorific value may be determined usingexternal data, such as a look-up table for a tracer substance in thefuel, or data encoded in a barcode associated with the fuel, or otherstored data.

The determination may therefore be performed in many different ways. Forexample:

-   -   a barcode of a fuel to be added to a fuel tank 50, 53 of the        aircraft 1 may be scanned to read data of the fuel, or a tracer        substance (e.g. a dye) identified and fuel properties looked up        based on that tracer;    -   data may be manually entered, or transmitted to the aircraft 1;    -   a fuel sample may be extracted for ground-side analysis prior to        take-off;    -   fuel properties may be inferred from measurements of the        propulsion system 2 activity during one or more periods of        aircraft operation, e.g. engine warm-up, taxi, take-off, climb        and/or cruise; and/or    -   one or more fuel properties may be detected onboard, optionally        in-flight, for example using in-line sensors and/or other        measurements.

The calorific value of the fuel may therefore be looked up, physicallydetermined from the results of combustion (either on-wing or off-wing),chemically determined from detected chemical properties,calculated/inferred from other detected fuel properties (either on-wingor off-wing), or otherwise determined or obtained.

In some examples, combinations of these techniques may be used todetermine and/or verify one or more fuel characteristics, includingcalorific value.

Fuel characteristics may be detected in various ways, both direct (e.g.from sensor data corresponding to the fuel characteristic in question)and indirect (e.g. by inference or calculation from othercharacteristics or measurements). The characteristics may be determinedas relative values as compared to another fuel, or as absolute values.For example, one or more of the following detection methods may be used:

-   -   The aromatic or cycloparaffin content of the fuel may be        determined based on measurements of the swell of a sensor        component made from a seal material such as a nitrile seal        material.    -   Trace substances or species, either present naturally in the        fuel or added to act as a tracer, may be used to determine fuel        characteristics such as the percentage of sustainable aviation        fuel in the fuel or whether the fuel is kerosene.    -   Measurements of the vibrational mode of a piezoelectric crystal        exposed to the fuel may be used as the basis for the        determination of various fuel characteristics including the        aromatic content of the fuel, the oxygen content of the fuel,        and the thermal stability or the coking level of the fuel—for        example by measuring the build-up of surface deposits on the        piezoelectric crystal which will result in a change in        vibrational mode.    -   Various fuel characteristics may be determined by collecting        performance parameters of the gas turbine engine 10 during a        first period of operation (such as during take-off) and then        during a second period of operation (e.g. during cruise)        comparing these collected parameters to expected values if using        fuel of known properties.    -   Various fuel characteristics including the aromatic hydrocarbon        content of the fuel may be determined based on sensor        measurements of the presence, absence, or degree of formation of        a contrail by the gas turbine 10 during its operation.    -   Fuel characteristics including the aromatic hydrocarbon content        may be determined based on a UV-Vis spectroscopy measurement        performed on the fuel.    -   Various fuel characteristics including the sulphur content,        naphthalene content, aromatic hydrogen content and hydrogen to        carbon ratio may be determined by measurement of substances        present in the exhaust gases emitted by the gas turbine engine        10 during its use.    -   Calorific value of the fuel may be determined in operation of        the aircraft 1 based on measurements taken as the fuel is being        burned—for example using fuel flow rate and shaft speed or        change in temperature across the combustor 16.    -   Various fuel characteristics may be determined by making an        operational change arranged to affect operation of the gas        turbine engine 10, sensing a response to the operational change;        and determining the one or more fuel characteristics of the fuel        based on the response to the operational change.    -   Various fuel characteristics may be determined in relation to        fuel characteristics of a first fuel by changing a fuel supplied        to the gas turbine engine 10 from the first fuel to a second        fuel, and determining the one or more fuel characteristics of        the second fuel based on a change in a relationship between T30        and one of T40 and T41 (the relationship being indicative of the        temperature rise across the combustor 16). The characteristics        may be determined as relative values as compared to the first        fuel, or as absolute values by reference to known values for the        first fuel.

As used herein and discussed below, T30, T40 and T41, and any othernumbered pressures and temperatures, are defined using the stationnumbering listed in standard SAE AS755, in particular:

-   -   T30=High Pressure Compressor (HPC) Outlet Temperature;    -   T40=Combustion Exit Temperature; and    -   T41=High Pressure Turbine (HPT) Rotor Entry Temperature.

In some examples, the obtaining the calorific value of the fuelavailable to the aircraft 1 and/or of the fuel already onboard theaircraft 1 prior to refuelling may comprise chemically and/or physicallydetermining the calorific value of the available fuel. Equivalentdeterminations may be performed for the fuel(s) to be used forrefuelling—either using ground-based testing (e.g. prior to refuelling)or onboard testing (e.g. using one or more sensors 204 on a fuel supplyline within the aircraft 1, leading to a tank 50, 53). Optionally, thatdetermination may be or comprise performing one or more of:

-   -   (i) identifying a tracer in the available fuel, such as a dye or        a marker trace element, and looking up a calorific value        corresponding to that tracer;    -   (ii) inferring the calorific value from one or more detected        physical and/or chemical properties of the available fuel, for        example using a ground-based testing-unit to analyse a fuel        sample; and/or    -   (iii) combusting a sample of the available fuel to determine its        calorific value directly, optionally using the gas turbine        engine 10.

The step of chemically and/or physically determining the calorific valueof the available fuel may therefore be performed on the aircraft 1 insome examples.

The inventors appreciated that, as different fuels can have differentproperties whilst still conforming to the standards, knowledge of thefuel(s) available to an aircraft 1 can allow more efficient, tailored,control of the propulsion system 2. For example, changing to a fuel witha higher calorific value may allow a smaller amount (mass and/or volume)of fuel to supply an aircraft's energy needs for a flight. As more poweris needed to lift a greater mass of fuel, taking sufficient fuel for theintended flight (including a safety margin above the expected energydemand), but not completely filling the tank(s) 50, 53, may provide anefficiency bonus by reducing take-off weight of the aircraft 1.Knowledge of the calorific value of the fuel can therefore be used as atool to improve aircraft performance, e.g. avoiding carrying excess fuelweight.

In the example shown in FIG. 4 , two sensors 204 a, 204 b are provided,each arranged to physically and/or chemically detect one or morefeatures of the composition of the fuel being added to the fuel tank 50,53 on refuelling. One or more of these fuel characteristics may be usedto infer or calculate calorific value. The sensors may be described aspart of a refuelling management system 204, 206.

In alternative examples, no such sensors 204 may be provided (forexample, a barcode associated with the fuel storage vessel 60 may beread and the corresponding data on the fuel provided to the refuellingmanager 206 rather than sensing fuel properties), or more or fewer,and/or differently-located, sensors may be provided.

In the examples described herein with respect to FIGS. 4 and 5 , priorto refuelling an aircraft 1, an amount of energy required for anintended flight profile of that aircraft is determined. As used herein,the “amount of energy required for an intended flight profile” generallyincludes a safety margin above a calculated amount of fuel needed forthe aircraft 1, with its expected take-off weight, to complete theintended flight. The safety margin may comprise an allowance forpossible diversion to an alternative destination airport, e.g. in caseof weather conditions preventing a safe landing at the intendeddestination airport.

As used herein, the term “flight profile” refers to the operationalcharacteristics (e.g. height/altitude, power setting, flight path angle,airspeed, and the like) of an aircraft 1 as it flies along a flighttrack, and also to the trajectory/flight track (route) itself. Indetermining the amount of energy required to complete an intended flightprofile, external conditions may also be considered (e.g. weather, as aprevailing headwind can increase the energy required, for example).Additionally or alternatively, the size of the safety margin may beselected to cover variation in external conditions.

When an aircraft 1 is about to be refuelled, in various examples, acalorific value of the fuel available to the aircraft 1 (i.e. the fuelwith which the aircraft is to be refuelled) is obtained, optionallyusing any of the off-wing methods described above. The total specificenergy of the fuel available (either per unit mass or per unit volume)can therefore be established. An amount of fuel required to meet theenergy requirements of the flight profile may then be determined. Insome cases, an on-wing determination method could be used (e.g.combusting a sample or use of onboard sensors). On-wing determinationmay lead to a need to pause and re-start refuelling whilst the desiredamount of fuel is calculated.

In many cases, the tank(s) 50, 53 of the aircraft 1 are not completelyempty when the aircraft 1 is to be refuelled. In such cases, thecalorific value of fuel already onboard the aircraft 1 is also obtained,optionally by retrieving characteristics of the fuel (e.g. calorificvalue, or other properties from which calorific value can be calculated)from storage, and/or by combusting a sample of the fuel (e.g. duringtaxiing) and measuring the calorific value directly. From a known amountof remaining fuel (mass, volume, or % full of the tank(s) 50, 53) andthe known calorific value, the total energy available from the onboardfuel can be determined.

The total energy available from the onboard fuel can therefore besubtracted from the determined amount of energy required for theintended flight profile to determine an amount of energy needed to beadded to the aircraft 1 on refuelling. Using the specific energy (orvolumetric energy density) of the fuel available, a desired amount (i.e.mass or volume) of the fuel available to the aircraft 1 can therefore becalculated.

In some examples, an aircraft 1 may have multiple tanks 50, 53 whichcontain different fuels, optionally with different calorific values.Total fuel energy content for onboard fuel may therefore be summedacross the different tanks 50, 53. Similarly, if an aircraft 1 is to berefuelled with multiple different fuels, the determined amount of energyrequired can be split between the different fuels as desired, and a massor volume of each fuel to be added can be calculated.

The aircraft 1 can then be refuelled with the calculated amount (mass orvolume) of the available fuel.

In some examples (especially examples in which data look-up has beenused to determine one or more calorific values), after refuelling acheck is then performed to verify the calorific value of the fuel. Thecheck may comprise measuring the calorific value of the fuel in use inthe gas turbine engine 10, optionally during warm-up of the engine priorto taxiing, or during taxiing of the aircraft 1 prior to take-off.Ideally, the fuel characteristics may be determined (and any checks of,or based on, the determined characteristics performed) before theaircraft 1 leaves the ground, in case the fuel is not as expected.

The method 2020 performed in these examples is illustrated in FIG. 5 .

At step 2022, an amount of energy required for an intended flightprofile, including a safety margin, is obtained. In this context, theamount of energy being “obtained” means that a numerical value ormeasure for the amount of energy required is in some way made availablefor use in the present method 2020—whether this is by calculation orother determination (e.g. from a look-up table), by communication withanother system, by input from a user (e.g. at a graphical userinterface), by retrieval from memory, or in any other suitable way. Forexample, the refuelling manager 206 may receive a message including avalue indicative of the energy requirement, or may calculate such avalue based on knowledge of the intended flight profile and of theaircraft 1 and aircraft load.

At step 2024, a calorific value of fuel available to the aircraft 1 isobtained. In this context, the calorific value being “obtained” againmeans that a numerical value or measure of the calorific value is insome way made available for use in the present method 2020. The valuemay be obtained in any suitable way, for example being manually entered,e.g. via a graphical user interface in communication with the refuellingmanager 206, electronically communicated to the refuelling manager 206,e.g. by wired or wireless communication from a barcode scanner followingreading of a barcode, and/or determined from sensor data. The value maybe stored for future use, optionally in memory of ,or in communicationwith, the refuelling manager 206.

The steps of obtaining 2022 the required amount of energy and obtaining2024 a calorific value of the fuel available for refuelling purposes maybe performed in either order, or simultaneously.

It will be appreciated that, if the tank 50, 53 is not empty prior torefuelling, the energy content of fuel already onboard may also beobtained—e.g. by retrieving a previously-stored calorific value for thatfuel, and calculating the energy content of the remaining fuel byreference to that and the amount of fuel remaining in the tank(s) 50,53, or by calculation from sensor data.

At step 2026, an amount of the available fuel needed to provide therequired energy is calculated, based on the obtained values. Thecalculated amount may be a volume and/or a mass of fuel—refuelling maybe performed volumetrically or gravimetrically. The calculation may beperformed by the refuelling manager 206 itself in some examples. Inexamples in which there is remaining fuel onboard, the energy of thatfuel may be subtracted from the energy required for the intended flightprofile before calculation of the amount of the available fuel needed.

At step 2028 the calculated amount (mass or volume) of the availablefuel needed is output so as to allow the aircraft 1 to be refuelledaccordingly. For example, a refuelling manager 206 may display theamount at a graphical user interface, may transmit the amount to anothersystem for the attention of a refuelling technician, or may provide thevalue to an automated system so as to cause that system to automaticallyterminate the refuelling process, or trigger an alert, once thecalculated amount of fuel has been provided.

Updated values for the calorific value of the fuel in the fuel tank 50,53 after refuelling may be calculated and stored for future use,optionally by the refuelling manager 206. In examples in which anaircraft 1 has multiple fuel tanks 50, 53 which are fluidly linked suchthat the fuels in the tanks 50, 53 are equivalent, a single calorificvalue for the fuel may be stored and updated. In examples in which anaircraft 1 has multiple fuel tanks 50, 53 which are not fluidly linked,such that there may be differences between fuels in the different tanks50, 53, a separate calorific value may be stored and updated for eachtank 50, 53.

The method 2020 may then be iterated on each refuelling event.

The method 2020 may further comprise controlling 2029 an automatedrefuelling system to refuel the aircraft 1 in line with the calculatedamount of fuel. Alternatively, the refuelling may be performed and/orterminated manually, e.g. by a person reading the output calculatedamount (mass or volume) of the available fuel needed, optionally from adisplay associated with the aircraft's fuel line connection port 62.

In some examples, refuelling may be started before one or more of thepreceding method steps 2022-2028 are performed, those steps may beperformed during refuelling, and the refuelling process may then beterminated when the calculated amount of fuel has been added.

In some examples, especially in implementations in which a calorificvalue of fuel is manually input at step 2024, a check may then beperformed to verify the input data. The check may comprise measuring thecalorific value of the fuel in use in the gas turbine engine 10,optionally during taxiing of the aircraft 1 prior to take-off. Thedetermining the calorific value of the fuel may be performed bymonitoring engine parameters during a first time period of aircraftoperation during which the gas turbine engine 10 uses the fuel; anddetermining the calorific value of the fuel based on the monitoredengine parameters (e.g rate of fuel combustion, fuel flow rate,temperatures, pressures, generated thrust, shaft speed, etc.). In caseof a significant mis-match in the entered and determined values, theaircraft 1 may return to the terminal for further investigation, as asafety measure.

A propulsion system 2 for an aircraft 1 may therefore comprise arefuelling manager 206 arranged to:

-   -   obtain 2022 an amount of energy required for an intended flight        profile, optionally including a safety margin;    -   obtain 2024 a calorific value of fuel available to the aircraft;    -   calculate 2026 the amount (mass or volume) of the available fuel        needed to provide the required energy; and    -   output 2028 the amount (mass or volume) of the available fuel        needed so as to allow the aircraft 1 to be refuelled        accordingly.

In some implementations, a refuelling manager may be providedseparately; not onboard the aircraft 1. Optionally, a refuelling managermay be provided as part of a ground-based refuelling station.

If the flight profile does not include a safety margin, or if anyreasons to extend a safety margin are found (e.g. forecast adverseweather or abnormally high load), a (further) safety margin may be addedat the calculation stage.

The refuelling manager 206 may additionally be arranged to control 2029an automated refuelling system such that it refuels the aircraft 1 inline with the calculated amount of fuel (e.g. adding the calculatedamount of fuel only, within tolerances).

The refuelling manager 206 may be provided as a separate refuellingmanagement unit built into the propulsion system 2, and/or as softwareand/or hardware incorporated into the pre-existing aircraft controlsystems (e.g. as part of an Engine Electronic Controller (EEC) 42). Insome examples, the refuelling manager 206 may store calorific value datafor any fuel currently in the aircraft's tank(s) 50, 53, and may causethose data to be updated following refuelling. The data may be storedseparately from circuitry performing the calculations and retrieved whenrequired—wherever the data are stored, that storage can be thought of asa part of the refuelling manager 206, whether or not it is integral orphysically connected in any way.

In examples in which a calorific value is looked up or read (e.g. fromscanning a barcode), this value may be automatically transferred to therefuelling manager 206, or may be typed into a user interface of, orassociated with, the refuelling manager 206, e.g. by a pilot ortechnician.

In some implementations, a non-transitory computer readable mediumhaving stored thereon instructions to cause the method 2020 describedabove may be provided, for use on-wing or off-wing. The instructions maycause the processor to:

-   -   obtain 2022 an amount of energy required for an intended flight        profile of an aircraft 1;    -   obtain 2024 a calorific value of fuel available to the aircraft;    -   calculate 2026 the amount of the available fuel needed to        provide the required energy; and    -   output 2028 the mass or volume of the available fuel needed so        as to allow the aircraft to be refuelled accordingly.

In implementations with automated refuelling, the instructions may befurther arranged to cause the processor to control 2029 fuel input tothe aircraft 1, such that the aircraft is refuelled with the calculatedamount of the available fuel. In other implementations, the output maybe read by a person who then refuels the aircraft 1 accordingly, and/orused to provide an alert when a sufficient volume or mass of fuel hasbeen added.

The inventors appreciated that, as different fuels can have differentproperties, whilst still conforming to the standards, knowledge of thefuel(s) available to an aircraft 1 can allow more efficient, tailored,control of the propulsion system 2. For example, changing to a fuel witha higher calorific value may allow a smaller amount of fuel to supply anaircraft's energy needs for a flight. As more power is needed to lift agreater mass of fuel, taking sufficient fuel for the intended flight(including a safety margin above the expected energy demand), but notcompletely filling the tank(s), may provide an efficiency bonus byreducing take-off weight of the aircraft. Knowledge of the calorificvalue of the fuel can therefore be used as a tool to improve aircraftperformance, e.g. avoiding carrying excess fuel weight and/or gaugingthe likelihood of forming contrails behind an aircraft in givenatmospheric conditions. The calorific value of one or more fuel(s) maybe checked by a second method to improve reliability, especially ifknowledge of the calorific value is to be used in any way that mightinfluence safety (e.g. ensuring sufficient fuel onboard for a safelanding at a destination airport).

In various examples as described with respect to FIGS. 6 to 9 , anaircraft 1 receives an input of calorific value data for fuel providedto the aircraft 1 on refuelling. For example, a barcode of a fuel to beadded to a fuel tank 50, 53 of the aircraft may be scanned to read dataof the fuel, or a tracer substance (e.g. a dye) identified and fuelproperties looked up based on that tracer. Alternatively, a pilot ortechnician may be provided with a calorific value, e.g. on a writtenlabel or orally, for entry into an aircraft system. Calorific value datamay be manually entered, or electronically transmitted to the aircraft1. In examples in which a calorific value is looked up or read (e.g.from scanning a barcode), this value may be automatically transferred toa fuel tracking system 203 of the aircraft 1, or may be typed into auser interface of, or associated with, the fuel tracking system 203 e.g.by a pilot or technician.

It will be appreciated that checking of calorific values of fuels may beimportant, especially in examples involving manual data entry ofcalorific values. The received input of the calorific value data forfuel provided to the aircraft 1 on refuelling may therefore be checked,and an alert may be provided if the determined calorific value of fuelis inconsistent with the calorific value data input received.

In the example shown in FIG. 6 , two sensors 204 a, 204 b are provided,each arranged to physically and/or chemically detect one or morefeatures of the composition of the fuel being added to the fuel tank 50,53 on refuelling. One or more of these fuel characteristics may be usedto infer or calculate calorific value. The sensors 204 may be describedas part of a fuel tracking system 203, which may be described as a fuelcomposition tracking system 203.

In alternative examples, no such sensors 204 may be provided (forexample, a barcode associated with a fuel storage vessel 60 may insteadbe read and the corresponding data on the fuel provided to therefuelling manager 206), or more or fewer, and/or differently-located,sensors may be provided.

In many cases, the tank(s) 50, 53 of the aircraft 1 are not completelyempty when the aircraft 1 is to be refuelled. In such cases, thecalorific value of fuel already onboard the aircraft 1 is also obtained,optionally by retrieving characteristics of the fuel (e.g. calorificvalue, or other properties from which calorific value can be calculated)from storage, and/or by combusting a sample of the fuel (e.g. duringtaxiing) and measuring the calorific value directly. From a known amountof remaining fuel (mass, volume, or % full of the tank(s)), a knownamount of added fuel, and the known calorific values of each, thecalorific value of the resultant blend can be determined whereappropriate (e.g. where fuels mix in the same tank or linked tanks, orare blended in use).

In various examples, the calorific value of fuel arranged to be suppliedto the gas turbine engine 10 in use is determined. It will beappreciated that, in implementations with multiple fuel sources, thefuel arranged to be supplied to the gas turbine engine 10 may differfrom the fuel in any given tank 50, 53. This fuel may comprise a mixtureof fuel with which the aircraft 1 has recently been refuelled, and anyfuel remaining in the aircraft's tank(s) 50, 53 following an earlierrefuelling event. In some examples, a sample may be extracted, e.g. froma pipe approaching the gas turbine engine 10, for analysis, for exampleusing a ground-based analysis unit or laboratory. In alternative oradditional examples, the determination may be performed on-wing, usingone or more sensors or other detectors onboard the aircraft 1.

In alternative or additional examples, the calorific value of the fuelprovided to the aircraft 1 on refuelling is determined—for example, byground-based testing prior to the refuelling event, or by on-wingtesting in a pipe or tank containing only the newly-added fuel. Inexamples with an active fuel management system, it may be possible tosupply the new fuel alone to the gas turbine engine 10, irrespective ofwhether or not older fuel is still on board (e.g. by tank/fuel sourceselection), and to calculate the calorific value of the new fueldirectly from engine performance.

In examples in which ground-based testing of the fuel is performed, thecalorific value of the fuel provided to the aircraft 1 on refuelling maybe determined in a fuel testing unit 61 a. In some examples, the fueltesting unit may form a part of a fuel delivery apparatus 61. In otherexamples, a fuel testing unit 61 a may be provided by an airline orairport, or a sample may be sent to a laboratory for analysis.

The testing unit 61 a of such examples may be provided off-wing at arefuelling site.

In examples in which on-wing testing of the fuel is performed, thedetermining the calorific value of the fuel supplied to the gas turbineengine 10 in use may be performed during at least one of taxiing of theaircraft prior to take-off, and during climb. If a sufficiently largediscrepancy is identified, the flight may be aborted subject tore-verification of the fuel.

In examples in which on-wing testing of the fuel is performed, thedetermining the calorific value of the fuel supplied to the gas turbineengine 10 in use may be performed by burning fuel taken from the fueltank 50, 53 in an auxiliary power unit of the aircraft 1 and measuringthe calorific value. Alternatively or additionally, the main gas turbineengine 10 itself may be used to perform the combustion.

After refuelling, a check is then performed to verify the calorificvalue of the fuel onboard the aircraft 1. The check may comprisemeasuring the calorific value of the fuel in use in the gas turbineengine 10, optionally during taxiing of the aircraft 1 prior totake-off.

If a discrepancy between the determined calorific value and the receivedcalorific value data exceeds a threshold, an alert may be provided—e.g.an audible and/or visual alarm, and/or a message sent to a pilot orother responsible party. It will be appreciated that, when thedetermined calorific value is that of a fuel mixture created on-boardthe aircraft 1, the determined value is not compared directly to thereceived calorific value data; instead, adjustments are made to one orthe other to reflect the contribution from the older fuel.

The method 2030 performed is illustrated in FIG. 7 .

At step 2032, an input of calorific value data for a fuel provided tothe aircraft 1 on refuelling is received. This data may be manuallyentered, transmitted to the fuel tracking system 203, obtained byreading a bar code, or similar. The input may comprise a calorific valuefor the fuel, and/or may be arranged to allow the calorific value forthe fuel to be determined (e.g. calculated or looked up).

The input may be obtained in any suitable way, for example beingmanually entered, e.g. via a graphical user interface in communicationwith the fuel tracking system 203, and/or electronically communicated tothe fuel tracking system 203, e.g. by wired or wireless communicationfrom a barcode scanner following reading of a barcode. The input may beor comprise sensor data in some implementations, e.g. providing one ormore chemical or physical fuel characteristics. A calorific value of thefuel in, or determined from, the input may be stored for future use,optionally in memory of, or in communication with, the fuel trackingsystem 203.

At step 2034, a calorific value is determined for at least one of:

-   -   (i) the fuel supplied to the gas turbine engine 10 in use (that        fuel including some of, or being, the fuel provided to the        aircraft 1 on refuelling); and    -   (ii) the fuel provided to the aircraft 1 on refuelling.

The determination 2034 does not use the calorific value data provided atstep 2032, such that the determinations of calorific values in steps2032 and 2034 are independent of each other. The determination 2034 mayotherwise use any approach described above. The value may be stored forfuture use, optionally in memory of, or in communication with, the fueltracking system 203.

It will be appreciated that the fuel supplied to the gas turbine engine10 in use may differ from the fuel provided to the aircraft 1 onrefuelling, as other fuel already on board the aircraft 1 may be used(either mixing in the same tank 50, 53 as the new fuel, or beingprovided in a blend from different onboard fuel sources).

Steps 2032 (receiving calorific value input) and 2034 (determiningcalorific value, independently of that input) may be performed in eitherorder, or simultaneously.

At step 2036, an alert is provided if the determined calorific value ofthe fuel is inconsistent with the calorific value data input received.

It will be appreciated that, if the tank 50, 53 is not empty prior torefuelling, the calorific value of fuel already onboard may also beobtained—e.g. by retrieving a previously-stored calorific value for thatfuel—and may also be used in determining 2034 whether or not thedetermined calorific value of the fuel is inconsistent with thecalorific value data input received.

Updated values for the calorific value of the fuel in the fuel tank 50,53 after refuelling may be calculated and stored for future use,optionally by the fuel tracking system 203. In examples in which anaircraft 1 has multiple fuel tanks 50, 53 which are fluidly linked suchthat the fuels in the tanks 50, 53 are equivalent, a single calorificvalue for the fuel may be stored and updated. In examples in which anaircraft 1 has multiple fuel tanks 50, 53 which are not fluidly linked,such that there may be differences between fuels in the different tanks50, 53, a separate calorific value may be stored and updated for eachtank.

The method 2030 may then be iterated on each refuelling event.

Implementations of the method 2030 therefore allow a check to beperformed to verify the input data. The check may comprise measuring thecalorific value of the fuel in use in the gas turbine engine 10,optionally during warm-up of the engine 10 or taxiing of the aircraft 10prior to take-off. The determining 2034 the calorific value of the fuelmay be performed by monitoring engine parameters during a first timeperiod of aircraft operation during which the gas turbine engine 10 usesthe fuel; and determining the calorific value of the fuel based on themonitored engine parameters (e.g rate of fuel combustion, temperatures,pressures, shaft speed(s) generated thrust, etc.). In case of asignificant mis-match in the entered and determined values, the aircraft1 may return to the terminal for further investigation, as a safetymeasure, and/or propulsion system control based on knowledge of fuelcalorific value may not be performed.

A propulsion system 2 for an aircraft may therefore comprise a fueltracking system 203 arranged to:

receive an input of calorific value data for fuel provided to theaircraft 1 on refuelling;

determine at least one of:

-   -   i. the calorific value of fuel supplied to the gas turbine        engine 10 in use; and    -   ii. the calorific value of the fuel provided to the aircraft 1        on refuelling; and provide an alert if the determined calorific        value is inconsistent with the calorific value data input        received.

The fuel tracking system 203 may be provided as a separate fuel trackingunit built into the propulsion system 2, and/or as software and/orhardware incorporated into the pre-existing aircraft control systems. Insome examples, the fuel tracking system 203 may store calorific valuedata for any fuel currently in the aircraft's tank(s) 50, 53, and maycause those data to be updated following refuelling. The data may bestored separately from circuitry performing the calculations andretrieved when required—wherever the data are stored, that storage canbe thought of as a part of the fuel tracking system 203, whether or notit is integral or physically connected in any way.

The fuel tracking system 203 comprises a fuel composition tracker 202.The fuel composition tracker 202 of the example being describedcomprises memory 202 a arranged to store the current fuel characteristicdata (in particular, calorific values), and processing circuitry 202 carranged to calculate updated values for the one or more fuelcharacteristics of the fuel in the fuel tank 50, 53 after refuelling.The calculated values may then replace the previously stored fuelcharacteristic data in the memory, and/or may be time- and/ordate-stamped and added to the memory. A log of fuel characteristic datawith time may therefore be assembled.

The fuel composition tracker 202 of the example shown also includes areceiver 202 b arranged to receive data relating to fuel composition(e.g. calorific value, or data which can be used to determine calorificvalue) and/or requests for fuel composition information.

The fuel composition tracker 202 of the example shown also includes anoutput module 202 c arranged to provide an alert if the determinedcalorific value of fuel supplied to the gas turbine engine 10 and/or ofthe newly-added fuel is inconsistent with the calorific value data inputreceived, for example by sending a signal to an alarm or other system.

In examples in which a calorific value is looked up or read (e.g. fromscanning a barcode), this value may be automatically transferred to thefuel composition tracker 202, or may be typed into a user interface of,or associated with, the fuel composition tracker 202, e.g. by a pilot ortechnician.

In some implementations, the calorific value checking method 2030 may beincorporated into a refuelling method 2020 as described earlier withrespect to FIG. 5 , as illustrated in FIG. 9 . In such implementations,the fuel composition tracker 202 may also be referred to as a refuellingmanager, or may form part of a refuelling manager 206.

At step 2022, an amount of energy required for an intended flightprofile, optionally including a safety margin, is obtained. In thiscontext, the amount of energy being “obtained” means that a numericalvalue or measure for the amount of energy required is in some way madeavailable for use in the present method 2020, 2030—whether this is bycalculation or other determination (e.g. from a look-up table), bycommunication with another system, by input from a user (e.g. at agraphical user interface), by retrieval from memory, or in any othersuitable way. For example, the refuelling manager 202, 206 may receive amessage including a value indicative of the energy requirement, or maycalculate such a value based on knowledge of the intended flight profileand of the aircraft 1 and aircraft load.

The term “flight profile” is used as defined above.

At step 2032, an input of calorific value data for a fuel provided tothe aircraft 1 on refuelling is received, as described above.

Steps 2022 and 2032 may be performed in either order, or simultaneously.

It will be appreciated that, if the tank 50, 53 is not currently empty,the energy content of fuel already on board may also be obtained—e.g. byretrieving a previously-stored calorific value for that fuel, andcalculating the energy content of the remaining fuel by reference tothat and the amount of fuel remaining in the tank(s) 50, 53.

At step 2026, an amount of the available fuel needed to provide therequired energy is calculated, based on the obtained values. Thecalculated amount may be a volume and/or a mass of fuel—refuelling maybe performed volumetrically or gravimetrically. The calculation may beperformed by the refuelling manager 206 itself in some examples. Inexamples in which there is remaining fuel onboard, the energy of thatfuel may be subtracted from the energy required for the intended flightprofile before calculation of the amount of the available fuel needed.

At step 2028 the calculated amount (mass or volume) of the availablefuel needed is output so as to allow the aircraft 1 to be refuelledaccordingly. For example, a refuelling manager 206 may display theamount at a graphical user interface, may transmit the amount to anothersystem for the attention of a refuelling technician, or may provide thevalue to an automated system so as to cause that system to automaticallyterminate the refuelling process once the calculated amount of fuel hasbeen provided.

Updated values for the calorific value of the fuel in the fuel tank 50,53 after refuelling may be calculated and stored for future use,optionally by the refuelling manager 202, 206. In examples in which anaircraft 1 has multiple fuel tanks 50, 53 which are fluidly linked suchthat the fuels in the tanks 50, 53 are equivalent, a single calorificvalue for the fuel may be stored and updated. In examples in which anaircraft 1 has multiple fuel tanks 50, 53 which are not fluidly linked,such that there may be differences between fuels in the different tanks50, 53, a separate calorific value may be stored and updated for eachtank.

The method 2020 may further comprise controlling 2029 an automatedrefuelling system to refuel the aircraft 1 in line with the calculatedamount of fuel. Alternatively, the refuelling 2029 may be performedmanually, e.g. by a person reading the output calculated amount (mass orvolume) of the available fuel needed, optionally from a displayassociated with the aircraft's fuel line connection port 62.

In some examples, refuelling may be started before one or more of steps2022 to 2028 are performed; those steps may be performed duringrefuelling, and the refuelling process may then be terminated when thecalculated amount of fuel has been added.

A check 2034, 2036 is then performed to verify the input data, asdescribed above.

The check may comprise measuring 2034 the calorific value of the fuel inuse in the gas turbine engine 10, optionally during engine warm-up ortaxiing of the aircraft 10 prior to take-off. The determining 2034 thecalorific value of the fuel may be performed by monitoring engineparameters during a first time period of aircraft operation during whichthe gas turbine engine 10 uses the fuel; and determining the calorificvalue of the fuel based on the monitored engine parameters (e.g. rate offuel combustion, temperatures, shaft speeds, pressures, generatedthrust, etc.). In case of a significant mis-match in the entered anddetermined values, an alert is provided 2036 and the aircraft 1 mayreturn to the terminal for further investigation, as a safety measure.

The check may be performed during or after refuelling.

The method 2020, 2030 may then be iterated on each refuelling event.

In some examples, the fuel tracking system 203 may therefore be furtherarranged to:

-   -   obtain 2022 an amount of energy required for an intended flight        profile, including safety margin;    -   obtain 2032 a calorific value of fuel available to the aircraft;    -   calculate 2036 the amount (mass or volume) of the available fuel        needed to provide the required energy; and    -   output 2038 the amount (mass or volume) of the available fuel        needed so as to allow the aircraft to be refuelled accordingly.

In such examples, the fuel tracking system 203 may be referred to as arefuelling manager 206. It will be appreciated that the examplesdescribed with respect to FIGS. 4 and 5 may therefore be combined withthe examples described with respect to FIGS. 6 to 9 .

In such examples, the calculating the mass or volume of the availablefuel needed to provide the required energy comprises obtaining acalorific value of fuel already in the fuel tank 50, 53, and subtractingthat from the determined amount of energy required for the intendedflight profile.

The method may further comprise calculating a calorific value for themixed fuel after refuelling. The calorific value of fuel supplied to thegas turbine engine 10 may only be determined to be inconsistent with thecalorific value data input received when the calculated calorific valuefor the mixed fuel does not match the determined calorific value of fuelsupplied to the gas turbine engine 10 in use.

The inventors appreciated that, as different fuels can have differentproperties, whilst still conforming to the standards, knowledge of thefuel(s) available to an aircraft 1 can allow more efficient, tailored,control of the propulsion system. For example, changing to a fuel with ahigher calorific value may allow a smaller flow rate of fuel to meet anaircraft's energy needs at a particular point in the flight envelope, sopotentially providing more fuel to auxiliary systems (e.g. fueldraulicactuators or fuel-oil heat exchangers) if a total pumped flow rate isconstant.

Knowledge of the calorific value of the fuel can therefore be used as atool to improve aircraft performance in flight even if it is not knownat the point of refuelling, e.g. gauging the likelihood of formingcontrails behind an aircraft 1 in given atmospheric conditions andchanging fuel source or altitude mid-flight. Further, in-flightdetermination of calorific value may be used as a check to verify dataprovided on or before refuelling.

As discussed above, calorific values (also referred to as heatingvalues) of fuels may be directly determined—for example by measuring theenergy released when a certain volume or mass of the fuel is combustedin the gas turbine engine 10—or calculated from other fuel parameters.In the examples presently being described, with respect to FIGS. 10 to12 , the calorific value of the fuel is directly determined by measuringthe energy released (or taking measurements which allow the energyreleased to be inferred) when a certain volume or mass of the fuel iscombusted within a gas turbine engine 10, 44 of the aircraft 1. Theperformance of the main/propulsive gas turbine engine 10 itself, and/orof a gas turbine engine of an Auxiliary Power Unit (APU) 44 is thereforeused to determine this fuel characteristic.

In addition to the propulsion system 2 described with respect to FIGS. 4and 6 , a power system 4 of the implementation shown in FIG. 10 includesan Auxiliary Power Unit (APU) 44. The more general term “power system” 4may be used in place of propulsion system 2 when the system 2,4 does notprovide only propulsive power, or indeed when the system 4 does notprovide any propulsive power. A propulsion system 2 is an example of apower system 4.

The APU 44 is a gas turbine engine smaller than those 10 on the wings ofthe aircraft 1, and is arranged to provide electrical power to systemsof the aircraft 1; for example, lighting, heating, air conditioningand/or similar. The APU 44 may be, for example, an APU in Honeywell's331 Series, such as the HGT 1700 auxiliary power unit (APU). In someimplementations, the APU 44 may be certified for in-flight use; in otherimplementations, it may be certified for ground use only. An aircraftAPU 44 is generally arranged to be started using one or more aircraftbatteries so as to provide electrical power as well as optionally bleedair for air conditioning and for engine start. The APU 44 of theimplementation shown is located towards the rear of the fuselage, and isnot arranged to provide any propulsive power to the aircraft 1. Inalternative implementations, the APU may be differently located (e.g.within a nacelle 21 of the aircraft 1), and/or may provide somepropulsive power. In the example shown in FIG. 10 , the centre fuel tank50 is arranged to supply fuel to the APU 44; fuelling arrangements mayvary in other examples.

Engine parameters during a first time period of aircraft operationduring which the gas turbine engine 10, 44 uses the fuel are sensed, andoptionally monitored over time. Those parameters may include one or moreof the thrust/propulsion provided, the volume (or mass) of fuel used ina given time (e.g. calculated from a fuel pumping rate/fuel flow rate,bearing in mind a pump spill ratio if applicable), the exhausttemperature, one or more shaft speeds, one or more temperature readingsof other components/in other locations, and/or one or more pressuremeasurements. In some examples, it may be assumed that 100% of the fuelfed to a combustor 16 of the gas turbine engine 10 is fully combusted.In other examples, different assumptions on the completeness ofcombustion may be made as applicable.

Sensors 224, for example temperature sensors 224 a and pressure sensors224 b, may be provided in association with the or each gas turbineengine 10 so as to monitor the performance of the gas turbine engine 10.

Based on the monitored engine parameters, a calorific value of the fuelcan then be determined.

As discussed above, it will be appreciated that checking of calorificvalues of fuels may be important, especially in examples involvingmanual data entry of calorific values. The determined calorific valuemay therefore be checked by a second determination (and optionally alsoadditional determinations), and/or used to check a manually entered (orotherwise provided) value, and an alert may be provided if thedetermined calorific value of the fuel is inconsistent with thecalorific value data input received. Examples as currently beingdescribed with respect to FIGS. 10 to 12 may therefore be combined withexamples described above with respect to FIGS. 6 to 9 .

In various examples, an aircraft 1 receives an input of calorific valuedata for fuel provided to the aircraft 1 on refuelling and a comparisonof the input value with the determined value may provide that check. Forexample, a barcode of a fuel to be added to a fuel tank 50, 53 of theaircraft 1 may be scanned to read data of the fuel, or a tracersubstance (e.g. a dye) identified and fuel properties looked up based onthat tracer. Alternatively or additionally, a pilot or technician may beprovided with a calorific value, e.g. on a written label or orally, forentry into an aircraft system. Calorific value data may therefore bemanually entered, or transmitted to the aircraft 1. In examples in whicha calorific value is looked up or read (e.g. from scanning a barcode),this value may be automatically transferred to a fuel tracking system203 of the aircraft 1, or may be typed into a user interface of, orassociated with, the fuel tracking system 203, e.g. by a pilot ortechnician.

In some implementations, the calorific value determined by measuring orotherwise determining the energy released when a certain volume or massof the fuel is combusted may be verified against a value determined in adifferent way, such as any of the determination methods described above.

The calorific value of the fuel may therefore be physically determinedfrom the results of combustion on-wing, and optionally verified withvalues provided to the aircraft 1, chemically determined from detectedchemical properties, or calculated/inferred from other detected fuelproperties (either on-wing or off-wing), for example using any of thedetection techniques described above. In alternative or additionalexamples, the process of determining calorific value based on engineperformance may be repeated, and the values compared, so providing acheck using the same approach. In alternative or additional examples, acalorific value based on engine performance may be determined in twodifferent engines, so providing a check; e.g. two different propulsivegas turbine engines 10, or a propulsive gas turbine engine 10 and an APUgas turbine engine 44.

Repeating the sensing and determination so as to obtain a secondcalorific value may be described as performing the same method in asecond time period. The first and second time periods may be atdifferent stages of aircraft operation (e.g. ground operations vs.cruise, climb vs. cruise, or cruise at a first altitude vs. cruise at asecond altitude), or may be at the same stage of aircraft operation -therefore, in some cases, no changes may be made to propulsion systemcontrol and/or the ambient conditions may be the same between the twotime periods, such that the same values of the monitored parameters(within errors/natural variation) would be expected.

If a discrepancy between the determined calorific value and the receivedcalorific value data exceeds a threshold, an alert may be provided—e.g.an audible and/or visual alarm, and/or a message sent to a pilot orother responsible party. It will be appreciated that, when thedetermined calorific value is that of a fuel mixture created onboard theaircraft 1, the determined value is not compared directly to thereceived calorific value data; instead, adjustments are made to one orthe other to reflect the contribution from the different fuel(s).

If a sufficiently large discrepancy is identified, the flight may beaborted subject to re-verification of the fuel, and/or no fuel-specificcontrol of the propulsion system 2 may be performed.

In some examples, an aircraft 1 may comprise multiple fuel tanks 50, 53,and two or more of the fuel tanks may contain different fuels, which mayhave different calorific values. In such cases, a determination of thecalorific value for the fuel in each tank 50, 53 may be made—for exampleusing a fuel management system to initially provide 100% fuel from onetank, and doing a first determination, and then using the fuelmanagement system to initially provide 100% fuel from the other tank,and doing a second determination. Alternatively or additionally,calorific values for two (or more) different blends of the two (or more)different fuels may be determined from the engine performance on thespecific blends, and calorific values for the fuels in the individualtanks 50, 53 may be calculated based on those determinations. In othersuch cases, the two or more different fuels may always be provided inthe same ratio in a blend, and the calorific value of the blend maytherefore be the only value of interest.

In the examples described herein, the calorific value of fuel suppliedto the gas turbine engine 10 in use is determined. This fuel maycomprise a mixture of fuel with which the aircraft 1 has recently beenrefuelled, and any fuel remaining in the aircraft's tank(s) 50, 53following an earlier refuelling event, and may comprise a mixture offuels from different tanks.

The determination(s) are performed in a first stage of aircraftoperation, and the results can then be used to influence control of thepropulsion system 2, or of the aircraft 1 more generally, in one or morelater stages of aircraft operation. For example, the determining thecalorific value of the fuel supplied to the gas turbine engine 10 in usemay be performed during taxiing of the aircraft 1 prior to take-off,and/or during other ground-based operations, and the result may be usedto influence control during one or more of take-off, climb, and cruise.Especially in implementations in which an Auxiliary Power Unit 44 isused for the determination, the determination may even be performed atan airport gate, before the main gas turbine engine(s) 10 have beenturned on. Alternatively, the determination may be performed undercruise conditions and the results used to influence control later in theflight, e.g. during cruise later in the same flight.

It will be appreciated that cruise conditions generally account for alarge proportion of most commercial flights, and that optimisingpropulsion system control for cruise therefore provides optimisation forthe bulk of the flight envelope in most cases.

The method 2040 performed is illustrated in FIG. 11 .

At step 2042, engine parameters are sensed or monitored during a firsttime period in which a fuel of interest is being combusted in a gasturbine engine 10, 44 of the aircraft 1. This time period may bedescribed as a first time period of aircraft operation, and may occur atany point of aircraft operation, including when the aircraft 1 isstationary (e.g. at a gate). Particularly in examples in which theengine parameters are monitored whilst the aircraft 1 is stationary, thegas turbine engine used may be an APU 44 of the aircraft 1, rather thanone of the engines 10 arranged primarily to provide propulsive power.

The first time period may be a period at idle at initial start of theengine. This may allow fuel calorific value to be determined before aflight commences, and operation of the engine 10 thereafter may bevaried according to the determined calorific value. The method 2040described herein may therefore be used as part of an active controlscheme.

By way of example of monitored engine parameters, the mass fuel flowrate into the combustor 16, shaft speed of one or more shafts 26, 27 ofthe engine 10, and/or one or more pressures and temperatures may besensed—either instantaneously at a point within the first time period,or with monitoring over the first time period. The fuel flow rate intothe combustor 16 may be measured directly—many current aircraft 1 have afuel flow meter at that location, and a meter could be added if not.Alternatively, the fuel flow rate on entry to the combustor may beinferred from data collected elsewhere—e.g. from the position of a fuelmetering valve (such valves commonly provide positional feedback), orfrom one or more pumps or flow meters located elsewhere. In exampleswith a volumetric pump rather than a gravimetric pump, mass flow ratemay be calculated from the volumetric flow rate and the fuel density, orthe calorific value calculation may be adjusted accordingly.Additionally, it will be appreciated that current aircraft 1 routinelymonitor one or more shaft speeds, and that this information, like otherdata, is often provided to an Engine Electronic Controller (EEC) 42.Step 2042 may therefore be performed without the need for any newsensors.

At step 2044, a calorific value is determined for the combusted fuel,using the monitored engine parameters.

For example, the mass fuel flow rate used to obtain a given speed of thelow pressure shaft, 26, or indeed the speed of any of the engine shafts26, 27, may be used to calculate the calorific value of the fuel (eitherdirectly or by using a look-up table of known values). In someimplementations, an instantaneous measurement may be performed so as todetermine a calorific value. In other cases, the parameters (e.g. fuelflow rate and shaft speed) may be monitored over a longer period ofoperation of the same fuel, for example for improving confidence of thecalculation.

The change in relationship between mass fuel flow rate and shaft speedmay be at least approximately 1:1 to the change in calorific value ofthe fuel (assuming no gearing of the shaft).

In some implementations, measurements may be taken in both the firsttime period in which the first fuel of interest is being combusted in agas turbine engine 10 and in a second time period in which a second fuelof interest is being combusted in the gas turbine engine 10. Other thanthe change in fuel, all other engine control options/all other featuresof engine operation may be kept constant between the first and secondtime periods, such that the fuel change is the only variable and aresponse of the system (in terms of a change to one or more of themonitored parameters) can be attributed to the change in fuel alone. Thechange in relationship between the mass fuel flow rate and the speed ofthe selected shaft may have a relationship close to 1:1 to the change incalorific value. The calorific value of the second fuel may therefore bedetermined based on knowledge of the calorific value of the first fuel,optionally as a relative value compared to that of the first fuel.

As an alternative to considering shaft speed, mass fuel flow may be heldconstant on changing fuel, and a change (if any) in the temperature riseacross the combustor 16 considered. The combustor exittemperature—T40—may be compared to the compressor exittemperature—T30—to get a measure of this temperature change (with thecompressor exit temperature corresponding closely to the combustor entrytemperature).

As used herein, T30 and T40, and any other numbered temperatures, aredefined using the station numbering listed in standard SAE AS755, inparticular:

-   -   T30=High Pressure Compressor (HPC) Outlet Temperature    -   T40=Combustion Exit Temperature

In current engines 10, T40 is generally not measured directly usingconventional measurement technology, such as thermocouples, due to thehigh temperature. A direct temperature measurement may be takenoptically but, alternatively or additionally, a value for T40 mayinstead be modelled or inferred from other measurements (e.g. usingreadings from thermocouples used for temperature measurement at otherstations and knowledge of the gas turbine engine architecture andthermal properties).

For a fuel with a higher calorific value, an increase in the temperaturerise across the combustor (T40-T30) would be expected, and vice versa,so mass fuel flow rate and combustor temperature change can be used asan alternative to shaft speed, or as a check of a calculation based onshaft speed.

For the examples listed, calculation of a change in calorific value onfuel change is described. It will be appreciated that absolute valuesmay be calculated, but that looking at a change in shaft speed and/orcombustor temperature rise (or another parameter) for a fixed mass flowrate of fuel may offer improved accuracy in cases in which there arerelevant accuracy limits on how well fuel mass flow can be measured.

When changes are assessed, as described above, it may be desirable tohave the first and second time periods as close together as reasonablypossible—a small interval may be left to ensure a complete change offuel in the combustor 16 and allow for any transient effects to pass.The required interval size (if any) may depend on fuel flow rate at theoperating condition. The gas turbine engine 10 generally reacts almostinstantly (within a second) to differences in fuel once that fuelreaches the combustor 16, and speed probes used for shaft speedmeasurements generally have a low time constant. At relatively lowpower, low fuel flow rate conditions, an interval of around ten secondsfrom when the fuel entering the pylon which connects the engine 10 tothe airframe of the aircraft 1 changes may be used. At higher power,where fuel flow rate may be four or more times higher, and interval of2-3 seconds from fuel change on pylon entry may be appropriate. It willbe appreciated that travel time from a fuel tank to the engine 10 mayvary based on tank location as well as fuel flow rate, and can beaccommodated accordingly with knowledge of the specific aircraft 1—pylonentry is therefore mentioned here for ease of generalisation, althoughtime change from opening or closing of a valve at or near a fuel tank50,53, or activation or deactivation of a fuel pump 108, may be used invarious implementations, with the interval calculated with reference tofuel flow time between the point of interest and the engine 10. Theinterval may therefore be arranged to allow for time taken to flush thefirst fuel out from the fuel supply pipes and for the second fuel toreach the combustor 16, as well as for a new steady state to be reached.

Further, measurements may be averaged over a period of time (e.g. 5seconds up to 30 seconds) within each time period, or in the second timeperiod only, and any trends examined, to check that a new steady statehas been reached and/or to improve reliability. In other examples, thetransient behaviour itself may be used in the determination—no intervalmay be left in such cases, and parameters may be monitored during asingle time period covering the change.

In some implementations, the propulsion system 2 of the aircraft 1 maythen be controlled 2046 differently, depending on the determinedcalorific value of the fuel.

The method 2040 may also be used to obtain near-instantaneousmeasurements in flight, for example checking fuel calorific value whenfuel is drawn from a different tank 50, 53 or combination of tanks. Thismethod 2040 may then be used as part of an active control scheme tocontrol subsequent operation of the gas turbine engine 10. A change incalorific value may be detected, e.g. based on a change in therelationship between fuel flow rate and shaft speed, and engineperformance may be controlled accordingly, for example by changing oneor more of fuel flow rate; pump spill; altitude; guide vane staging(where variable-position vanes are provided); and fuel (where multipledifferent fuels are available onboard the aircraft 1).

As such, once the calorific value(s) of one or more fuels onboard theaircraft 1 have been determined, the propulsion system 2 can becontrolled based on the determined calorific value(s) by, for example:

-   -   Changing an operating parameter of a heat management system of        the aircraft 1 (e.g. a fuel-oil heat exchanger), or changing the        temperature of fuel supplied to the combustor 16 of the engine        10.    -   When more than one fuel is stored aboard an aircraft 1,        selecting which fuel to use for which operations (e.g. for        ground-based operations as opposed to flight, or for operations        with different thrust demands) may be made based on the fuel. A        fuel delivery system may therefore be controlled appropriately        based on the fuel characteristics.    -   Adjusting one or more flight control surfaces of the aircraft 1        so as to change route and/or altitude based on knowledge of the        calorific value of the fuel(s).    -   Changing the spill percentage of a fuel pump 108 (i.e. the        proportion of pumped fuel recirculated instead of being passed        to the combustor 16, as discussed in more detail below). The        pump 108 and/or one or more valves may therefore be controlled        appropriately based on the fuel characteristics.    -   Changing the scheduling of variable-inlet guide vanes (VIGVs).        The VIGVs may therefore be moved, or a movement of the VIGVs be        cancelled, as appropriate based on the fuel characteristics.

It will be appreciated that fuel flow rate on entry to the combustor 16is generally already measured in modern gas turbine engines 10, with agravimetric fuel flow meter often being provided. Conversions may bemade as applicable where a volumetric fuel flow meter is provided.Further, fuel flow rate may additionally or alternatively be inferredfrom position of a fuel metering valve which controls fluid flow intothe combustor 16 and/or other circuits—such fuel metering valvesgenerally provide positional feedback, but, especially at cruise asopposed to idle, improved accuracy may be provided by a flow meterdownstream of the fuel metering function. Similarly, shaft speeds arealready recorded in modern gas turbine engines 10, so no additionalhardware/sensors may be required to implement the method 2040 describedherein. The method 2040 may therefore be implemented in software,optionally as part of the EEC 42, without requiring any physical changesto the gas turbine engine 10.

A propulsion system 2, or other power system 4 as discussed above, foran aircraft 1 may therefore comprise a fuel tracker 202 arranged to:

-   -   monitor 2042 engine parameters during a first time period of        aircraft operation during which a gas turbine engine 10, 44 uses        the fuel; and    -   determine 2044 a calorific value of the fuel based on the        monitored engine parameters.

The fuel tracker 202 may then provide the determined calorific value asan output. The determined calorific value may be supplied to an aircraftcontrol system 42, for example to be used to influence control 2046 ofthe propulsion system 2.

The fuel tracker 202 may be provided as a separate fuel tracking unitbuilt into the propulsion system 2, and/or as software and/or hardwareincorporated into the pre-existing aircraft control systems. In someexamples, the fuel tracker may store calorific value data for any fuelcurrently in the aircraft's tank(s) 50, 53, and may cause those data tobe updated following new determinations (e.g. triggered by refuelling).The data may be stored separately from circuitry performing thecalculations and retrieved when required—wherever the data are stored,that storage can be thought of as a part of the fuel tracker 202,whether or not it is integral or physically connected in any way.

The fuel tracker 202 may form part of a fuel tracking system 203.

The fuel tracking system 203 comprises the fuel composition tracker 202.The fuel composition tracker 202 of the example being describedcomprises memory 202 a arranged to store received values for monitoredengine parameters and determined fuel calorific values, and processingcircuitry 202 c arranged to calculate calorific values based on thereceived values for monitored engine parameters. The calculatedcalorific value may replace previously stored fuel characteristic datain the memory, and/or may be time- and/or date-stamped and added to thememory. A log of fuel characteristic data (in particular, calorificvalue, although other characteristics may also be stored) with time maytherefore be assembled.

The fuel tracking system 203 comprises, or is in communication with, oneor more sensors 224. The sensors 224, for example temperature sensors224 a and pressure sensors 224 b, are associated with a gas turbineengine 10 so as to monitor the performance of the gas turbine engine 10.Data from these sensors are used, optionally in conjunction with otherdata provided, to calculate fuel calorific value based on engineperformance when burning that fuel.

The fuel composition tracker 202 of the example shown also includes areceiver 202 b arranged to receive data relating to fuel composition(including the monitored engine parameters, or values calculatedelsewhere therefrom) and/or requests for fuel composition information.

The fuel composition tracker 202 of the example shown also includes anoutput module 202 d. In some implementations, the output module 202 dmay be arranged to provide an alert if the determined calorific value offuel supplied to the gas turbine engine 10 is inconsistent with acalorific value data input received, for example by sending a signal toan alarm or other system.

In some examples, the fuel tracking system 203 may therefore be furtherarranged to:

-   -   receive an input of calorific value data for fuel in an aircraft        1, or provided to the aircraft 1 on refuelling;    -   compare the input calorific value data to the determined        calorific value; and    -   provide an alert if the determined calorific value of fuel        supplied to the gas turbine engine 10 is inconsistent with the        calorific value data input received.

In alternative or additional implementations, the output module 202 dmay send a message—for example comprising a fuel calorific value, or acontrol instruction based on the calorific value—to an aircraft controlsystem 42, e.g. an Engine Electronic Controller (EEC), so as toinfluence control of the propulsion system 2 based on the fuelcharacteristics.

The inventors appreciated that, as different fuels can have differentproperties, whilst still conforming to the standards, knowledge of thefuel(s) available to an aircraft 1 can allow more efficient, tailored,control of the propulsion system 2, and in particular it may beappropriate to change spill rates around a fuel pump 108 of the gasturbine engine 10. For example, changing to a fuel with a highercalorific value may allow a smaller flow rate of fuel to the combustor16 to meet an aircraft's energy needs at a particular point in theflight envelope, so potentially providing more fuel to auxiliary systems(e.g. fueldraulic actuators or fuel-oil heat exchangers) if a totalpumped flow rate is constant.

As depicted in FIGS. 4, 6 and 10 and described above, an aircraft 1 maycomprise multiple fuel tanks 50, 53; for example a larger, primary fueltank 50 located in the aircraft fuselage, and a smaller fuel tank 53 a,53 b located in each wing. In other examples, an aircraft 1 may haveonly a single fuel tank 50. Many different fuel tank arrangements arepossible as described above, and the tanks 50, 53 may form a single fuelsource, or multiple fuel sources.

The fuel used in a gas turbine engine 10 of the aircraft may thereforevary during flight (where an aircraft 1 has multiple distinct fuelsources) as well as between flights (as an aircraft 1 may be refuelledwith a different fuel). The tanks 50, 53 may contain different fuels—forexample with one tank 50 containing a kerosene jet fuel and another tank53 containing a SAF, or a kerosene-SAF blend. The different fuels may bemixed en route to the combustor 16. The percentage of SAF in a fuelsupplied to the engine 10 may therefore vary between 0% and 100% duringoperation of the aircraft 1 in some examples. The SAF may have a density(ρ) of between 90% and 98% of the density of kerosene. The SAF may havea calorific value, (CV) of between 101% and 105% the calorific value ofkerosene (calorific value being as defined above). For example, thecalorific value of kerosene may be 43.1 MJ/kg (with a current minimum CVallowed in the fuel specification of 42.8 MJ/kg), whereas the calorificvalue of SAF may be 44.2 MJ/kg. The calorific value and density of fuelblends may vary accordingly with a density of 90-100% of that ofkerosene and a calorific value of 100% to 105% of that of kerosene.

The aircraft 1 of the examples currently being described comprises afuel pump 108 arranged to pump fuel from the one or more tanks 50, 53 tothe gas turbine engine 10. The fuel pump 108 has an inlet 108 a arrangedto receive fuel and an outlet 108 b via which fuel leaves the pump 108.The received fuel may be from a single tank, or may be a blend from acombination of tanks 50, 53. The fuel fed to the gas turbine engine 10may therefore have a different composition from any fuel stored in atank in some examples. The fuel may pass through one or more elements ofa heat management unit, or other engine components, en route to theinlet 108 a of the pump 108. Further, not all of the fuel leaving thepump 108 is supplied to the combustor 16; some is instead recirculated(“spilled”), and recirculated fuel generally forms a proportion of thefuel entering the pump inlet 108 a.

The proportion of fuel passing through the pump 108 which isrecirculated is referred to as the spill or spill percentage, i.e.:

${Spill} = \frac{\begin{matrix}{{{Total}{fuel}{flow}},Q,{{{at}{pump}{outlet}} -}} \\{{Fuel}{flow}{into}{combustor}}\end{matrix}}{{{Total}{fuel}{flow}},Q,{{at}{pump}{outlet}}}$

The recirculated fuel may include fuel returned directly from the pumpoutlet 108 b to the pump inlet 108 a, for example via a fuel return line46 (FIG. 13 , flow S₁). In other examples, no such fuel return line 46may be present. The recirculated fuel may include fuel diverted awayfrom the combustor entry to serve other roles in auxiliary systems orengine components (FIG. 13 , flow S₂), for example with the fuel actingas a heat transport medium in one or more heat exchangers, or as aworking fluid in one or more fueldraulic actuators. Further, some of therecirculated fuel may be returned to a fuel tank 50, 53 before lateruse. Spilled fuel (FIG. 13 , flows S₁ and S₂) may therefore be used toperform engine functions as well as to allow a pump 108 to keepoperating at a set flow rate even when there are fluctuations incombustor fuel demand. As used herein, “spill” therefore includes fuelused for any purpose other than being fed to the combustor 16, not justfuel sent straight from the pump outlet 108 b to the pump inlet 108 a orback to a tank 50, 53.

When an aircraft 1 is operating on kerosene at cruise, spills of 70-85%are common, with spill often reaching 98% at flight idle. Precise spilllevels depend on one or more of aircraft and engine designs, thrustdemand, ambient temperature, altitude, and stage of cruise (e.g. due toa higher aircraft weight due to additional fuel, and often a loweraltitude, at the start of cruise, and a lighter, and often higher,aircraft at the end of cruise). A minimum spill of at least 5% or 10%may be set to ensure sufficient fuel is provided to auxiliary systems45. The pump 108 may be sized for the Maximum Take Off (MTO) thrust ofan aircraft 1, optionally at low altitude. Around 10% of total pumpcapacity may generally be accounted for supplying auxiliary systems 45at MTO. At cruise as compared to MTO, a pump 108 with a rotation speedlinked to an engine shaft will still be spinning very quickly, but theflow demanded by the combustor 16 will be significantly lower than atMTO; the fraction of spilled excess flow is therefore generally muchhigher at cruise than at MTO.

A new pump 108 on any engine 10 will generally have some extra capacity(effectively over-supplying fuel) as its performance is expected todeteriorate over time—the % of total flow spilled may therefore begreater for newer pumps 108 than for older pumps. It will be appreciatedthat the total fuel offtake requirements for auxiliary systems 45 etc.are a function of how many auxiliary systems are being supplied withfuel, what the requirements of these systems 45 are in terms of flow,and when the flow is required (e.g. not all fueldraulic actuatorsgenerally move at the same time). A spare capacity margin is generallyalso provided such that more spilled fuel is generally available thanneeded.

One or more sensors may be provided to detect spill directly. Forexample, the fuel flow rate into the combustor 16 may be measureddirectly—many current aircraft 1 have a fuel flow meter at thatlocation, and a meter could be added if not—or inferred frommeasurements elsewhere, and subtracted from a known pump outlet flowvalue, Q, to provide a measure of the recirculated fuel flow.

The inventors appreciated that, as different fuels can have differentproperties whilst still conforming to the standards, use of SAF, orkerosene-SAF blends, may change the desired spill at given conditions.In particular, as SAF content increases, a desired spill may generallyreduce. One or more valves associated with the pump 108, with the fuelreturn line 46, and/or with the fuel-using auxiliary systems orcomponents 45 may be used to control the spill. The pump 108 and valves(not pictured) and the fuel supply lines connecting them together form afuel supply system 230, as illustrated in FIG. 13 and FIG. 14 .

A fuel-change spill ratio, R_(s), is defined as:

$R_{s} = \frac{{spill}{percentage}{at}{cruise}{using}{kerosene}}{{spill}{percentage}{at}{cruise}{using}a{fuel}{with}X\%{SAF}}$

R_(s) is therefore equal to 1 when X=0, i.e. when the fuel is a purekerosene fuel, but varies when X increases. It will be appreciated that,in calculating R_(s), conditions are taken to be the same other than thechange of fuel—i.e. same engine 10, same stage of flight, same altitude,etc.

In various examples, X % is at least 30% (X≥30), and R_(s) is greaterthan or equal to 1.003.

In various examples, when the fuel blend is 50% SAF by weight (X=50),the fuel-change spill ratio is at least 1.0066, and when X=100 such thatthe fuel is pure SAF, the fuel-change spill ratio is at least 1.0138.

The following relationship may apply in various examples where X % is atleast 10%, and optionally at least 30%, illustrating the relationshipbetween % SAF and spill:

${R_{s} \geq {1 + \frac{X}{10000}}},{X \geq {10}}$

The fuel supply system 230 is arranged to supply fuel to the combustor16 at an energy flow rate, C, optionally measured in Mega Watts, MW.This energy flow rate, C, may be controlled to meet/equal the combustorenergy demand to obtain a given power output/thrust in given conditions.Defining the density of kerosene as ρ_(K), the calorific value ofkerosene as CV_(K) and the density and calorific value of the fuelsupplied to the combustor 16 as ρ_(F) and CV_(F), respectively, it hasbeen found advantageous to control the pump flow rate, Q, and spill ratebased on fuel properties and combustor energy demand such that:

$R_{s} = {\frac{{spill}{percentage}{at}{cruise}{using}{kerosene}}{{spill}{percentage}{at}{cruise}{using}{the}{fuel}{with}X\%{SAF}} = \frac{Q - \frac{C}{\left( {{CV}_{K} \times \rho_{K}} \right)}}{Q - \frac{C}{\left( {{CV}_{F} \times \rho_{F}} \right)}}}$

In various examples, the gas turbine engine 10 is arranged such that,for an engine 10 with a maximum take-off thrust in the range from 400 kNto 500 kN, at cruise:

$\frac{Q - \frac{\left( {Q - 4595} \right)}{\left( {5.88 \times 10^{- 6} \times {CV}_{K} \times \rho_{K}} \right)}}{Q - \frac{\left( {Q - 4995} \right)}{\left( {5.88 \times 10^{- 6} \times {CV}_{F} \times \rho_{F}} \right)}} \leq R_{s} \leq \frac{Q - \frac{\left( {Q - 4995} \right)}{\left( {5.88 \times 10^{- 6} \times {CV}_{K} \times \rho_{K}} \right)}}{Q - \frac{\left( {Q - 4595} \right)}{\left( {5.88 \times 10^{- 6} \times {CV}_{F} \times \rho_{F}} \right)}}$

where Q is measured in Imperial Gallons per hour, CV in CHU/lb, and p inlb per Imperial Gallon.

Converting units to measure Q in litres/second, CV in MJ/kg, and ρ in kgper litre gives:

$\frac{Q - \frac{\left( {Q - 5.8} \right)}{\left( {0.014 \times {CV}_{K} \times \rho_{K}} \right)}}{Q - \frac{\left( {Q - 6.31} \right)}{\left( {0.014 \times {CV}_{F} \times \rho_{F}} \right)}} \leq R_{s} \leq \frac{Q - \frac{\left( {Q - 6.31} \right)}{\left( {0.014 \times {CV}_{K} \times \rho_{K}} \right)}}{Q - \frac{\left( {Q - 5.8} \right)}{\left( {0.014 \times {CV}_{F} \times \rho_{F}} \right)}}$

In various examples, the gas turbine engine 10 is arranged such that,for an engine 10 with a maximum take-off thrust in the range from 300 kNto 350 kN, the following relationship holds at cruise:

$\frac{Q - \frac{\left( {Q - 3457} \right)}{\left( {1.13 \times 10^{- 5} \times {CV}_{K} \times \rho_{K}} \right)}}{Q - \frac{\left( {Q - 3857} \right)}{\left( {1.13 \times 10^{- 5} \times {CV}_{F} \times \rho_{F}} \right)}} \leq R_{s} \leq \frac{Q - \frac{\left( {Q - 3857} \right)}{\left( {1.13 \times 10^{- 5} \times {CV}_{K} \times \rho_{K}} \right)}}{Q - \frac{\left( {Q - 3457} \right)}{\left( {1.13 \times 10^{- 5} \times {CV}_{F} \times \rho_{F}} \right)}}$

where Q is measured in Imperial Gallons per hour, CV in CHU, and ρ inpounds per Imperial Gallon.

Converting units to measure Q in litres/second, CV in MJ/kg, and ρ in kgper litre gives:

$\frac{Q - \frac{\left( {Q - 4.37} \right)}{\left( {0.027 \times {CV}_{K} \times \rho_{K}} \right)}}{Q - \frac{\left( {Q - 4.87} \right)}{\left( {0.027 \times {CV}_{F} \times \rho_{F}} \right)}} \leq R_{s} \leq \frac{Q - \frac{\left( {Q - 4.87} \right)}{\left( {0.027 \times {CV}_{K} \times \rho_{K}} \right)}}{Q - \frac{\left( {Q - 4.37} \right)}{\left( {0.027 \times {CV}_{F} \times \rho_{F}} \right)}}$

The fuel-change spill ratio, R_(s), may therefore be controlled based onthe pump output flow and fuel properties.

As mentioned above, a desired spill may vary based on one or more ofambient temperature, altitude, and stage of cruise. R_(s) may thereforebe varied accordingly.

In various examples, R_(s) is decreased by less than 0.15%, andoptionally by more than 0.1%, between the beginning and end of cruise,for a constant temperature and altitude.

In various examples, R_(s) is decreased by at least 0.11% when altitudeincreases by at least 600 m.

A method 3000 performed in some embodiments is illustrated in FIG. 15 .

At step 3002, a fuel from one or more of the fuel tanks 50, 53 issupplied to the gas turbine engine 10. The fuel supplied to the gasturbine engine 10 comprises X % SAF, where X % is in the range from 5%to 100%, with any remainder of the fuel being kerosene. The fuel'sdensity is denoted as ρ_(F), and its calorific value as CV_(F).Correspondingly, ρ_(K) and CV_(K) are used for kerosene.

In some implementations, at step 3004, the propulsion system 2 iscontrolled such that:

-   -   the fuel-change spill ratio, R_(s), of:

$R_{s} = \frac{{spill}{percentage}{at}{cruise}{using}{kerosene}}{{spill}{percentage}{at}{cruise}{using}{the}{fuel}{with}{}X\%{SAF}}$

is equal to:

$\frac{Q - \frac{C}{\left( {{CV}_{K} \times \rho_{K}} \right)}}{Q - \frac{C}{\left( {{CV}_{F} \times \rho_{F}} \right)}}$

Where Q is the fuel flow rate at the pump outlet, and C is the energyflow rate of fuel entering the combustor.

The fuel-change spill ratio, Rs, may therefore be controlled based onknowledge of the fuel and of current engine operation. The calorificvalue and/or density of the fuel may be determined or otherwise obtainedusing any one or more of the approaches described above.

The control 3004 may be iteratively repeated and updated if/when thefuel changes.

In alternative or additional implementations, and when the fuel suppliedto the combustor 16 at step 3002 comprises at least 30% SAF (i.e. X≥30)at step 3004, the propulsion system 2 is controlled such that:

$R_{s} = \frac{{spill}{percentage}{at}{cruise}{using}{kerosene}}{{spill}{percentage}{at}{cruise}{using}a{fuel}{with}X\%{SAF}}$

is greater than or equal to 1.003.

In various such implementations, spill is controlled such that:

$R_{s} \geq {{1 + \frac{x}{10000}}.}$

For example, when X is 50, the fuel-change spill ratio may be at least1.0066, and when X is 100, such that the fuel is pure SAF, and thefuel-change spill ratio may be at least 1.0138.

In various implementations of the method 3000, the gas turbine engine 10may be arranged such that R_(s)≤1.04, and/or the gas turbine engine 10is arranged such that R_(s)≥1.003, and optionally R_(s)≥1.014.

In other implementations, in particular in implementations in which acalorific value of the fuel is not known, the method 3000 may beflipped—instead of controlling R_(s) based on engine activity and fuelproperties, R_(s) may be iteratively adjusted on changing to a new fueluntil a desired energy flow rate, C, to the combustor 16 is achieved fora known (e.g. set or sensed) pump outlet flow rate Q. The change inspill, as captured in the ratio R_(s), may therefore be used todetermine the calorific value of a new fuel, if the fuel's density isknown, or a value of the calorific value multiplied by the fuel densityif not. A change in R_(s) may therefore be used to determine fuelproperties.

A propulsion system 2 for an aircraft 1 according to the examplescurrently being described may therefore comprise a gas turbine engine 10and one or more fuel tanks 50, 53 arranged to contain fuel to supplyfuel to power the gas turbine engine 10, one or more of the tanks 50, 53containing sustainable aviation fuel—SAF—either alone or as part of ablend. The SAF has a density of between 90% and 98% of the density, PK,of kerosene and a calorific value of between 101% and 105% the calorificvalue CVK, of kerosene.

The gas turbine engine 10 of many examples comprises an engine core 11comprising a turbine 19, a combustor 16, a compressor 14, and a coreshaft 26 connecting the turbine to the compressor; and a fan 23 locatedupstream of the engine core, the fan comprising a plurality of fanblades and being arranged to be driven by an output from the core shaft26; in addition to the fuel pump 108.

The fuel pump 108 is arranged to supply a fuel from one or more of thefuel tanks 50, 53 to the gas turbine engine 10 and provides a pumpoutput volumetric flow rate, Q. The fuel pump 108 has an inlet 108 aarranged to receive fuel from the one or more fuel tanks 50, 53 and anoutlet 108 b arranged to provide fuel to the gas turbine engine 10, andis arranged to recirculate (spill) excess fuel back from the outlet tothe inlet (directly or indirectly), the percentage of fuel passingthrough the pump which is recirculated being referred to as a spillpercentage. The fuel supplied to the gas turbine engine 10 comprises X %SAF, where X % is in the range from 5% to 100%, and is optionally atleast 30%, with any remainder of the fuel being kerosene. The fuelsupplied to the gas turbine engine 10 has a density, ρ_(F), and acalorific value CV_(F).

The propulsion system 2 is arranged such that the fuel-change spillratio, R_(s), is as described above.

The inventors have also appreciated that a measurement of the mass andvolume of the fuel F being used by the aircraft can be used for thedetermination of the characteristics of the fuel. In one example, such ameasurement of a fuel mass and volume can be performed during arefuelling process in which fuel is loaded onto the aircraft.

FIG. 16 illustrates an aircraft 1 which is connected to a fuel storagevessel 60 for refuelling as described above. The fuel storage vessel 60may be carried by a fuel supply vehicle (e.g. a fuel tanker) or may be afixed storage vessel from which the aircraft 1 can be refuelled. Theaircraft 1 comprises a fuel line connection port 62 which is coupled toa fuel loading line 61 during refuelling. The fuel loading line 61 maycomprise a fuel pipe of known design. The fuel line connection port 62is fluidly coupled with the fuel tanks 53, 55 of the aircraft 1 by afuel transmission line or lines 63 on board the aircraft so that fuelreceived via the fuel loading line 61 is transferred and stored withinthe fuel tanks 53, 55. The fuel loading line 61 and fuel transmissionline 63 may together form a fuel supply line used to supply fuel to thefuel tanks 53, 55 on board the aircraft 1 from the fuel storage vessel60. In some examples, the fuel transmission line(s) 63 may not bepresent, with the fuel instead delivered directly from a fuel lineconnection port for each fuel tank (or set of interconnected fueltanks).

Referring again to FIG. 16 , the aircraft 1 further comprises a fuelcharacteristic determination system 102. The fuel characteristicdetermination system 102 is arranged to determine one or more fuelcharacteristics of the fuel being loaded, or which has been loaded, ontothe aircraft 1, those characteristics being any of those described orclaimed herein.

The fuel characteristic determination system 102 generally comprises amass sensor 103, a volume sensor 104 and a fuel characteristicdetermination module 105. The mass sensor 103 is arranged to measure amass of fuel being loaded onto the aircraft 1. In the presentlydescribed example, the mass sensor 103 is arranged to measure a massflow rate of fuel as it flows from the fuel line connection port 62 tothe fuel tanks 55, 53 on board the aircraft 1. The mass sensor in thisexample may be a mass flow meter arranged to measure the mass of fluidtravelling past a fixed point within the fuel supply line per unit time.The point at which the mass flow rate is measured may be at any pointupstream of the aircraft fuel tanks 53, 55 in which the fuel is storedso that the mass flow rate of fuel being loaded on the aircraft 1 can bemeasured. In the present example, the mass flow sensor 103 is located ina fuel conduit (e.g. part of the fuel transmission line 63) fluidlyconnecting the fuel line connection port 62 to one of the aircraft fueltanks 53, 55. It may however be located at other points within theaircraft fuel system, for example at the fuel connection port 62. Themass flow meter 103 may be a Coriolis flow meter of known design. Anyother suitable type of mass flow meter may however be used, for exampleany in which the mass determination is not measured indirectly based onknowledge of the density of the fuel.

The volume sensor 104 in the presently described example is arranged tomeasure a volume flow rate of fuel being loaded onto the aircraft 1. Inthe presently described example, the volume sensor 104 is arranged tomeasure a volume flow rate of fuel as it flows from the fuel lineconnection port 62 to the fuel tanks 53, 55 on board the aircraft 1. Thevolume sensor 104 in this example may be a volume flow meter arranged tomeasure the volume of fluid travelling past a fixed point within theaircraft fuel system per unit time. Similarly to the mass flow meter103, the volume flow meter 104 may be located at any point upstream ofthe aircraft fuel tank(s) 53, 55 in which the fuel is stored. It maytherefore be provided in the same fuel conduit as the mass flow meter103, and may be downstream of the mass flow meter 103 as shown in thefigures, or upstream from it. Any suitable volumetric flow meter may beused, such as a turbine or pressure flow meter. The volume flow metermay be of a type which does not infer the volumetric flow rate from ameasured mass flow rate. Similarly, the mass flow meter may be of a typewhich does not infer the mass flow rate from a measured volumetric flowrate. The mass flow rate and volume flow rates are therefore measuredindependently from each other (and without requiring knowledge of thefuel density).

The fuel characteristic determination module 105 is arranged todetermine one or more fuel characteristics of the fuel being loaded ontothe aircraft 1 based on the fuel mass and fuel volume determined by themass sensor 103 and volume sensor 104. The fuel characteristicdetermination module 105 is therefore in communication with the mass andvolume sensors 103, 104 as shown in the Figures such that it can receivesignals from them which are indicative of the fuel mass and fuel volume.In the present example, the fuel characteristic determination module 105is a separate unit, and may be in communication with an electronicengine controller (EEC) 42 of each of the gas turbine engines 10provided on the aircraft 1. The determined one or more fuelcharacteristics may be communicated to the EEC 42 such that therespective engine 10 can be controlled accordingly as will be discussedbelow. In other examples, the determination module 105 may be part ofthe EEC 42 of the (or each) engine 10.

The fuel characteristic determined by the fuel characteristicdetermination module 105 may be any of those described or claimedherein. In order to determine a fuel characteristic, the fuelcharacteristic determination module 105 is arranged to compare the fuelmass and fuel volume and determine a corresponding fuel density (e.g. bydividing the mass flow rate by the volume flow rate). As fuels havingdifferent fuel characteristics will exhibit a known variation indensity, a characteristic of the fuel being loaded on to the aircraft 1can be inferred based on the density. In some embodiments, a fuelcharacteristic or characteristics may be determined by calculating adeviation from the density value that would be expected if the fuel werefossil kerosene. In other embodiments, the determination module 105 maybe arranged to access a lookup table defining known fuel characteristicdependence on fuel density. The measured fuel density can then becompared to the values in the lookup tables so that a characteristic ofthe fuel can be determined.

In one example, the fuel characteristic may be the percentage of SAF inthe fuel. The inventors have observed that SAF has a lower densitycompared to fossil kerosene, and this difference can be used to inferthe percentage of SAF present in the fuel based on a measurement of thedensity of the fuel as it is loaded onto the aircraft. Other fuelcharacteristics may also have an associated variation in fuel density.For example, the determination module 105 may determine that the fuel isfossil kerosene (e.g. substantially 100% fossil kerosene) if themeasured fuel density is that associated with a fossil kerosene fuelwith no SAF present. In other examples, the fuel characteristicsdetermined may include a hydrocarbon distribution of the fuel, or anaromatic hydrocarbon content of the fuel. In other examples, the densityof the fuel determined from the mass and volume measurements may beconsidered a fuel characteristic determined by the fuel characteristicdetermination system 102.

In the example shown in FIG. 16 , a mass flow rate and volume flow rateof fuel are measured in fuel flowing within a conduit leading to thefuel tanks on board the aircraft. In some examples, the mass sensor isarranged to measure the mass of fuel stored in the aircraft fuel tanksduring loading, or after it has been loaded. For example, the masssensor 103 may be arranged to measure an increase in weight of theaircraft 1 as fuel is loaded, or an increase in weight of the fuel tanks53, 55 themselves as they are filled with fuel. The fuel characteristicdetermination module 105 may in such examples base the fuelcharacteristic on a total mass of fuel loaded for a flight, or the massper unit time loaded onto the aircraft, associated with a measuredchange in weight. The same may be the case for the volume sensor 104. Insome examples therefore, the volume sensor 104 may be arranged todetermine a volume of fuel stored in the fuel tanks of the aircraft. Forexample, the volume sensor 104 may comprise one or more level sensorsarranged to measure a level of fuel within the, or each, fuel tank ofthe aircraft. The stored volume of fuel can then be determined. Thedetermination module 105 may in such examples base the fuelcharacteristic(s) on a total volume, or volume per unit time, of fuelloaded onto the aircraft. The determination module may therefore moregenerally receive signal indicative of the mass and/or volume of fuel(e.g. a weight or level measurement), rather than a direct mass orvolume measurement.

In various examples therefore, determining the mass of the fuel maycomprise measuring the mass and/or a change in the mass of any one ormore of: the aircraft; one or more fuel tanks on board the aircraft; afuel tanker vehicle from which the fuel is supplied; or a storage vesselfrom which the fuel is supplied to the aircraft. Determining the volumeof the fuel may comprise measuring the volume and/or a change in thevolume of fuel: (a) stored in one or more fuel tanks on board theaircraft; and/or (b) stored in a fuel storage vessel from which the fuelis supplied to the aircraft.

In the example illustrated in FIG. 16 , the determination system 102 islocated entirely on board the aircraft 1. In other examples, that maynot be the case, with one or more components of the determination system102 not being located on the aircraft 1. For example, the mass andvolume sensors 103, 104 and the determination module 105 may be includedin a dedicated unit which is separate from the aircraft 1. In someexamples, the fuel characteristic determination system 102 may belocated entirely outside of the aircraft. In such an example, the fuelcharacteristic determination module 105 may determine a fuelcharacteristic which is then communicated to the aircraft 1 (e.g. to acontrol module of the engines or engines 10). In this example, a datatransfer link may be provided (e.g. a wireless or wired data connection)and may be used to communicate the fuel characteristics to the aircraftfrom the fuel characteristic determination system 102. In some examples,the data transfer may be done manually by a user, e.g. a technician orother operator of the system may obtain the fuel characteristics fromthe determination system 102 and manually provide them to a controlmodule on board the aircraft.

In some examples, some of the components of the system 102 may belocated at the fuel supply vessel 60 (e.g. on board a fuel tankervehicle). For example, the sensors 103, 104 may be located within thefuel supply line 61, and the fuel characteristic determination module105 located on board the aircraft 1. In such an example, mass and volumemeasurements may be communicated to the determination unit on board theaircraft using any suitable (wired or wireless) data connection. In someexamples, the mass and volume sensors 103, 104 may be arranged tomeasure the mass and/or volume of fuel that is taken out of the storagevessel 60 connected to the aircraft 1. In one example, the mass sensor103 may be arranged to determine the mass of fuel supplied to theaircraft by measuring a change in weight of the fuel storage vessel 60,or a vehicle on which the fuel storage vessel 60 is transported.Similarly, the volume sensor 104 may comprise a level sensor arranged tomeasure the level of fuel contained in the fuel storage vessel 60.Measurements carried out on or associated with the fuel storage vesselmay be used to determine a total mass and/or total volume, or a mass perunit time and/or volume per unit time of fuel provided to the aircraft.

In some examples, the fuel characteristic determination module 105 isarranged to determine the one or more fuel characteristics based atleast in part on a fuel temperature. This may allow variation in thedensity of the fuel caused by changes in the fuel temperature to bedistinguished from those associated with the fuel characteristic orcharacteristics being determined. In some examples, the determinationmodule 105 is arranged to obtain a signal indicative of a currenttemperature of the fuel, for example from a fuel temperature sensorarranged to directly measure the temperature of the fuel, or an ambienttemperature sensor.

In the presently described examples, the fuel characteristicdetermination module 105 is arranged determine the one or more fuelcharacteristics based only on the determined fuel mass and fuel volume.In other examples, the determination module 105 may be arranged tocombine the fuel mass and fuel volume information with inputs from othersensors or other methods of determining fuel characteristics asdescribed elsewhere herein. This may allow a greater range of types offuel characteristic to be inferred, or may improve the accuracy of thefuel characteristic determination.

FIG. 17 illustrates a method 1002 of determining one or more fuelcharacteristics of an aviation fuel that may be carried out by the fuelcharacteristic determination system 102 shown in FIG. 16 . The method1002 comprises determining 1003 a mass of fuel being loaded (or whichhas been loaded) onto the aircraft 1; determining 1004 a fuel volume ofthe fuel being loaded (or which has been loaded) onto the aircraft 1;and determining 1005 one or more fuel characteristics based on thedetermined mass and volume of the fuel. The mass and volume measurementsmay be performed during a refuelling process e.g. to measure a massand/or volume of fuel as it is being loaded onto the aircraft, or thetotal mass and/or volume of fuel that has been loaded onto the aircraftonce the refuelling is complete (e.g. before operation of the aircraft).

Determining 1003 the fuel mass may comprise measuring a mass flow rate apoint within a fuel supply line used to convey fuel to the fuel tanks53, 55 of the aircraft. As discussed above, the point at which the massflow rate is measured may be any point upstream of one or more fueltanks 53, 55 on board the aircraft 1. For example, the fuel flow ratemay be measured at a point on board the aircraft through which fuelloaded onto the aircraft travels to reach the fuel tanks(s). In otherembodiments, the fuel flow rate may be measured at a point within a fuelloading system (i.e. not on the aircraft) such as a point within a fuelloading line 61 connected to the aircraft.

Determining 1004 the fuel volume may comprise measuring a volume flowrate a point within a fuel supply line used to convey fuel to the fueltanks 53, 55 of the aircraft. As discussed above, the point at which thevolume flow rate is measured may be any point upstream of one or morefuel tanks 53, 55 on board the aircraft 1 similarly to the measurementof the mass flow rate. The volume flow rate can be measured at similar,or an adjacent, position to the mass flow rate.

As discussed above, in some examples, determining 1003 the fuel mass maycomprise measuring the mass or a change in the mass of the aircraft 1,one or more fuel tanks 53, 55 on board the aircraft to which fuel issupplied, a tanker vehicle from which the fuel is supplied, or a fuelstorage vessel 60 from which it is provided. The change in weight of theaircraft 1, fuel tanks 53, 55, tanker vehicle, or fuel storage vessel 60may be used to determine a mass flow per unit time, or total mass, offuel loaded onto the aircraft 1 for comparison with a correspondingvolume measurement.

The step of determining 1004 the fuel volume may comprise measuring thevolume or a change in volume of fuel within one or more fuel tanks 53,55 on board the aircraft 1 or the fuel storage vessel 60. Measuring thechange in volume may comprise measuring a fluid level within therespective vessel/fuel tank in which the fuel is contained. The changein fuel level may be used to determine a volume per unit time, or totalvolume, of fuel loaded onto the aircraft 1 for comparison with thecorresponding mass measurement.

Determining 1005 the one or more fuel characteristics may comprisecomparing the determined fuel mass and fuel volume. As discussed above,this may include calculating 1006 a fuel density based on the mass andvolume. The one or more fuel characteristics may be determined 1005based on a comparison of the calculated fuel density with a knowndensity value associated with fuel having known characteristics. The oneor more fuel characteristics determined may be any of those describedherein which are associated with a corresponding characteristic fueldensity.

In the embodiment shown in FIG. 17 , the one or more fuelcharacteristics are further determined 1007 based on a signal indicativeof the temperature of the fuel. As discussed above, the signalindicative of the temperature of the fuel may be from a sensor arrangedto directly measure the fuel temperature, or a sensor arranged tomeasure the ambient temperature, or otherwise input to the determinationmodule.

In the present example, characteristics of fuel as it is being loadedonto the aircraft may be determined. In such examples, the one or morefuel characteristics determined may be communicated to the EEC 42directly if it is running during refuelling, or may otherwise be storedand communicated to the EEC when it is activated. If the EEC is notactive when fuel characteristics are determined, they may becommunicated to another control system of the aircraft.

Where fuel characteristics are determined for fuel being loaded onto theaircraft that fuel may be mixed with fuel already present in the fueltanks (e.g. from previous flights). The determined fuel characteristicsmay therefore be combined with those determined from previous times atwhich the aircraft was refuelled in order to determine thecharacteristics of the fuel stored in the aircraft fuel tanks. This maybe done using a summing method in which the amount of fuel loaded intothe tanks, the amount of fuel used during each flight, and thecorresponding characteristics of the fuel loaded are logged and combinedto determine the fuel characteristics of fuel actually stored within theaircraft tanks at a given time.

The inventors have also appreciated that the fuel characteristics can bedetermined during operation of the gas turbine engine, rather thanduring refuelling. FIG. 18 illustrates another example of a fuelcharacteristic determination system 106. In this example, the fuelcharacteristic determination system 106 is configured to determine oneor more characteristics of the fuel being used by the gas turbine engineof the present application during operation of the engine 10. Theexample of FIG. 18 therefore differs from that of FIG. 16 in that themass and volume of fuel that is delivered to a combustor of the engineis determined, e.g. the mass and volume may be measured while the fuelis burnt by the gas turbine engine 10. More specifically, this may beduring flight of the aircraft 1 to which the gas turbine engines 10 aremounted, or during operation of the aircraft 1 whilst it is on theground (e.g. during taxi).

FIG. 18 shows a schematic view of part of the fuel system of theaircraft 65 and combustion equipment 16 of the gas turbine engine 10.The combustion equipment 16 comprises a plurality of fuel nozzles (notshown in FIG. 18 ) arranged to inject fuel into a combustion can. Fuelis provided to the combustion equipment 16 by a fuel delivery regulator107 under the control of the EEC 42. Fuel is delivered to the fueldelivery regulator 107 by a fuel pump 108 from a fuel source 109 onboard the aircraft 1 (e.g. the fuel tanks 53, 55 described above). Thefuel delivery regulator 107 and combustion equipment 16 (which may bereferred to simply as a combustor) may be of known design, and may bearranged for staged (lean-burn) combustion or rich-burn combustion.

The fuel characteristic determination system 106 shown in FIG. 18generally comprises a fuel characteristic determination module 110, masssensor 111 and volume sensor 112. The system 106 shown in FIG. 18differs from that of FIG. 16 in that it is arranged to measure the massand volume of fuel as it is being supplied and burnt by the combustionequipment 16 of the gas turbine engine 10 as the engine is operating,rather than when the fuel is being loaded onto the aircraft 1.

The aircraft 1 comprises an aircraft fuel supply system located on boardthe aircraft which is suitable for suppling fuel F to each of the gasturbine engines 10 to be burnt in the engine combustion equipment 16 asdescribed above. The aircraft fuel supply system is arranged to providefuel to an engine fuel system provided on each of the gas turbineengines 10. The engine fuel system and aircraft fuel supply systemtogether form the (overall) fuel system of the aircraft 1 in which fuelis stored, delivered to the engine, and combusted. The fuel system ofthe aircraft includes any component which may store fuel, or throughwhich fuel flows during use or during refuelling.

The mass sensor 111 is arranged to measure a mass flow rate of fluidbeing supplied to the combustion equipment 16. In the presentlydescribed example, the mass flow sensor 111 is arranged to measure themass of fuel per unit time flowing between the fuel delivery regulator107 and the combustion equipment 16. Any suitable mass flow rate metermay be used for the mass flow sensor 111, such as a Coriolis flow meter.The mass flow meter may be any mass flow meter which performs a massmeasurement which is not based on knowledge of the density of the fuel.The mass flow sensor 111 may be arranged to measure the mass flow rateof fuel at any point within the fuel system of the aircraft that isupstream of the combustion equipment 16 (e.g. upstream of the fuelnozzles of the combustion equipment 16) and downstream of the fuelsource 109 from which the fuel is supplied on board the aircraft 1 (e.g.downstream of the one or more fuel tanks 53, 55 forming the fuelsource). In some examples, the mass flow rate is therefore measured at apoint within the engine fuel system such as in a fuel conduit within orforming part of the gas turbine engine 10 (rather than being measured bya sensor provided on the aircraft 1 to which the gas turbine engine 10is mounted). In some examples, the mass flow rate is measured at a pointimmediately before fuel the fuel is combusted, e.g. immediately beforeit enters the combustor. In yet other embodiments, the mass flow rate ismeasured at a point within the aircraft fuel supply system e.g. beforeit enters part of the gas turbine engine 10.

The volume sensor 112 is arranged to measure a volume flow rate of fluidbeing supplied to the combustion equipment 16. In the presentlydescribed example, the volume flow sensor is arranged to measure thevolume of fuel per unit time flowing between the fuel delivery regulator107 and the combustion equipment 16. Any suitable volume flow rate metermay be used for the volume flow sensor 112, such as a pressure orturbine type flow meter. The volume flow rate meter may be of a typethat is arranged to measure the volume flow rate without reliance on amass measurement. Similarly the mass flow rate meter may be of a typethat is arranged to measure the mass flow rate without reliance on avolume measurement. The volume and mass flow sensors may be of the sametype as described in the embodiment of FIG. 16 .

The volume flow sensor 112 may be arranged to measure the volume flowrate of fuel at any point within the fuel system of the aircraft that isupstream of the combustion equipment 16 (e.g. upstream of the fuelnozzles of the combustion equipment 16) and downstream of the fuelsource 109 from which the fuel is supplied (e.g. downstream of the oneor more fuel tanks 53, 55 forming the fuel source). It may therefore beat a similar or adjacent position to the mass sensor 111 (e.g. upstreamor downstream of it). In some examples, the volume flow rate istherefore also measured at a point within a fuel conduit within orforming part of the gas turbine engine 10 (rather than being measured bya sensor provided on the aircraft 1 to which the gas turbine engine 10is mounted). In some examples, the volume flow rate is measured at apoint immediately before fuel is combusted similarly to the mass flowrate, e.g. immediately before entering the combustor 16. In someexamples, the volume flow rate is measured in part of the aircraft fuelsupply system, before fuel reaches the gas turbine engines.

The fuel characteristic determination module 110 is in communicationwith the mass and volume sensors 111, 112, and is arranged to receivesignals therefrom indicative of the mass and volume of fuel per unittime being burnt by the combustion equipment 16. The fuel characteristicdetermination module 110 is arranged to determine one or more fuelcharacteristics of the fuel based on the fuel mass and volume similarlyto the determination module 105 described in connection with FIG. 16 .For example, the determination module 110 may be arranged to calculate adensity of the fuel from signal indicative of the fuel mass and volumebased on which a fuel characteristic can be found as described above. Ascan be seen in FIG. 18 , the determination module 110 may transmit theone or more fuel characteristics to the EEC 42. In other examples, itmay form part of the EEC 42.

The fuel characteristic(s) determined by the determination module 110may be any of those described or claimed herein. In order to determine afuel characteristic, the fuel characteristic determination module 110 isarranged to compare the fuel mass and fuel volume and determine acorresponding fuel density (e.g. by dividing the mass flow rate by thevolume flow rate). As fuels having different fuel characteristics willexhibit a known variation in density, a characteristic of the fuel beingloaded on to the aircraft 1 can be inferred. In some embodiments, thefuel characteristic may be determined by calculating a deviation fromthe density value that would be expected if the fuel were fossilkerosene. In other embodiments, the fuel characteristic determinationmodule 110 may be arranged to access a lookup table defining known fuelcharacteristic dependence on fuel density. The measured fuel density canthen be compared to the values in the lookup tables so that acharacteristic of the fuel can be determined. The fuel characteristicdetermination module may therefore operate in a corresponding manner tothat of FIG. 16 .

In one example, the fuel characteristic may be the percentage of SAF inthe fuel. The inventors have observed that SAF has a lower densitycompared to fossil kerosene, and this difference can be used to inferthe percentage of SAF present in the fuel based on a measurement of thedensity of the fuel as it is loaded onto the aircraft. Other fuelcharacteristics may also have an associated variation in fuel density.For example, the fuel characteristic determination module 110 maydetermine that the fuel is fossil kerosene (e.g. substantially 100%fossil kerosene) if the measured fuel density is that associated with afossil kerosene fuel with no SAF present. In other examples, the fuelcharacteristics determined may include a hydrocarbon distribution of thefuel, or an aromatic hydrocarbon content of the fuel. In other examples,the density of the fuel may be considered a fuel characteristicdetermined by the fuel characteristic determination system 106.

In the example illustrated in FIG. 18 , the mass sensor 111 and volumesensor 112 are arranged to measure the flow of fuel passing through theengine fuel system. In other embodiments, the determination module 110may be arranged to receive a signal indicative of the mass and/or volumeflow rate, which may not be a direct measurement of the flow of fuel.Such a signal may be an operating parameter of the fuel pump 108 such asthe pump speed or spill rate which has a known relationship with thefuel mass and/or volume flow rate.

In some examples, the fuel characteristic determination module 110 isarranged to determine the one or more fuel characteristics based on afuel temperature similarly to as described above. This may allowvariation in the density of the fuel caused by changes in the fueltemperature to be distinguished from those associated with the fuelcharacteristic or characteristics being determined. This may beimportant in examples where the fuel is heated above ambient temperatureby a heat exchanger forming part of a heat management system of theengine 10. In some examples therefore, the fuel characteristicdetermination module 110 is arranged to obtain a signal indicative of acurrent temperature of the fuel at the point where the volume and/ormass is measured. The temperature may be measured by a fuel temperaturesensor arranged to directly measure the temperature of the fuel, or anambient temperature sensor. In other examples, the fuel characteristicdetermination module 110 may receive a signal indicative of the fueltemperature which is inferred from other engine operating parametersrather than being based on a direct measurement.

In the presently described examples, the fuel characteristicdetermination module 110 is arranged to determine the one or more fuelcharacteristics based only on the determined fuel mass and fuel volumeobtained during operation of the gas turbine engine 10. In otherexamples, the fuel characteristic determination module 110 may bearranged to combine the fuel mass and fuel volume information withinputs from other sensors or other methods of determining fuelcharacteristics as described elsewhere herein. This may allow a greaterrange or types of fuel characteristic to be inferred, or improve theaccuracy of the fuel characteristic determination.

FIG. 19 illustrates a method 1008 of determining a fuel characteristicof an aviation fuel that may be carried out by the fuel characteristicdetermination system 106 shown in FIG. 18 . The method 1008 comprises:determining 1009 a mass of fuel being supplied to the combustor;determining 1010 a corresponding volume of the fuel being supplied tothe combustor; and determining 1011 one or more fuel characteristicsbased on the determined mass and volume. The mass and volume of fuel aredetermined during operation of the gas turbine engine.

As discussed above, the determining 1009 of the mass of fuel comprisesdetermining a mass flow rate of fuel being supplied to the combustor 16.The mass flow rate may be determined based on a measurement performed bya mass flow meter 111 on fuel flowing to the combustor 16. The mass flowrate may be measured at any point within the aircraft fuel system thatis upstream of the combustion equipment 16. For example, upstream of thefuel nozzles of the combustion equipment 16, and downstream of the fuelsource 109 from which the fuel is supplied on board the aircraft 1 (e.g.downstream of the one or more fuel tanks 53, 55 forming the fuelsource). In some examples, the mass flow rate may be measured at a pointin a fuel conduit of the engine fuel system. The mass flow rate may bemeasured immediately before fuel reaches the combustor 16.

Determining 1010 the volume of fuel comprises determining a volume flowrate of fuel being supplied to the combustor 16. The volume flow ratemay be determined 1010 based on a measurement of fuel flow using avolume flow meter 112 as described above. The volume fuel meter 112 maybe arranged to measure the volume flow rate at positions equivalent tothose described above with reference to the mass flow meter.

In some examples the mass and volume flow rate may be measured atpositions close to each other or immediately up or down stream from eachother, or at separate locations within the fuel supply system. In any ofthe examples herein, the mass and volume are measured for the same flowof fuel i.e. the mass and volume are measured at positions on a flowpath carrying the same fuel composition, and there are no additionalsources of sinks of fuel between. This allows the measured flow rates tobe meaningfully compared. In examples where fuels of differentcompositions are stored in the fuel tanks 53, 55 of the aircraft, themass and volume sensors are located so as to measure the same fuel flowe.g. both volume and mass sensors may be located in a pipe conveyingfuel from tank 53, which may store fuel of a different composition fromthat of tank 55. In such an example, a further mass and volume sensorpair may be provided in a pipe conveying fuel from tank 55 so that thecharacteristics of each fuel may be measured independently.

In some examples, the steps of determining 1009, 1010 the mass of fueland volume of fuel may comprise basing the determination on a signalfrom which the flow rate can be inferred rather than a directmeasurement. In some examples therefore, the determining of the mass orvolume of fuel is based on an operating parameter of the fuel pump 108or other signal indicative of the mass/volume flow rate.

As discussed above, determining 1011 the one or more fuelcharacteristics comprises comparing the determined fuel mass and fuelvolume. This may include calculating 1012 a fuel density based on themass and volume. The one or more fuel characteristics may be determined1011 based on a comparison of the calculated fuel density with a knownvalue associated with fuel having known characteristics. The one or morefuel characteristics determined may be any of those described hereinwhich are associated with a corresponding characteristic fuel density.

In the embodiment shown in FIG. 19 , the one or more fuelcharacteristics are further determined 1013 based on a signal indicativeof the temperature of the fuel. As discussed above, the signalindicative of the temperature of the fuel may be from a sensor arrangedto directly measure the fuel temperature, or a sensor arranged tomeasure the ambient temperature, or otherwise input to the determinationmodule from another source.

The fuel characteristics determined using any of the fuel characteristicdetermination systems or methods of determining a fuel characteristic inthe examples herein may be used in the operation of the aircraft, andmore specifically operation of the gas turbine engine(s) of theaircraft. This may allow the operation of the aircraft 1 to be modifiedin response to the fuel characteristic determined.

The present application therefore further provides a method 1065 ofoperating an aircraft 1 powered by one or more gas turbine engines 10 asillustrated in FIG. 20 . The method 1065 may be a method of operatingthe aircraft 1 of any of the examples described herein. The method 1065comprises determining 1066 one or more fuel characteristics. This maycomprise using any of the methods described herein. The method 1065further comprises operating 1067 the aircraft 1 according to the one ormore fuel characteristics. Operating the aircraft 1067 may morespecifically comprise operating the gas turbine engine(s) 10 mounted tothe aircraft 1, but may include operating other parts of the aircraft.

Once one or more fuel characteristics are known, the gas turbine engine10 or the aircraft more generally may be controlled or operated invarious different ways to take advantage of that knowledge. The step ofoperating 1067 the gas turbine engine or the aircraft may comprisemodifying 1067 a a control parameter of the aircraft, and specifically acontrol parameter of the gas turbine engine, in response to the one ormore fuel characteristics. Modifying the control parameter may includeany one or more of the following:

-   -   i) Modifying a control parameter of a heat management system of        the gas turbine engine (e.g. a fuel-oil heat exchanger) based on        the one or more fuel characteristics. By modifying the operation        of the heat exchanger the temperature of fuel supplied to the        combustor 16 of the engine 10 can be changed. In one example,        modifying the operation of the heat management system or        changing the temperature of the fuel may comprise increasing the        temperature of the fuel if the fuel characteristics indicate        that the fuel can tolerate operating at a higher temperature        without risk of coking or thermal breakdown.    -   ii) When more than one fuel is stored aboard an aircraft 1,        modifying a control parameter that controls a selection of which        fuel to use for which operations (e.g. for ground-based        operations as opposed to flight, for low-temperature start-up,        or for operations with different thrust demands) based on fuel        characteristics such as % SAF, nVPM generation potential,        viscosity, and calorific value. A fuel delivery system of the        aircraft may therefore be controlled appropriately based on the        fuel characteristics. The fuel delivery system may be controlled        to supply the engine with fuel having a different fuel        characteristic to that measured in step 1066. This may include,        for example, providing fuel with a relatively lower aromatic        content; providing fuel with a lower SAF content; or providing        fossil Kerosene fuel. The fuel supply may be controlled by        switching between fuel tanks, or changing a fuel blend ratio.    -   iii) Modifying a control parameter to adjust one or more flight        control surfaces of the aircraft 1, so as to change route and/or        altitude based on knowledge of the fuel.    -   iv) Modifying a control parameter to modify the spill percentage        of a fuel pump (i.e. the proportion of pumped fuel recirculated        instead of being passed to the combustor) of a fuel system of        the aircraft according to the one or more fuel characteristics,        for example based on the % SAF of the fuel. The pump and/or one        or more valves may therefore be controlled appropriately based        on the fuel characteristics.    -   v) Modifying a control parameter to change the scheduling of        variable-inlet guide vanes (VIGVs) based on fuel        characteristics. The VIGVs may be moved, or a movement of the        VIGVs be cancelled, as appropriate based on the fuel        characteristics.

In the examples above, the gas turbine engine or the aircraft isoperated according to the one or more fuel characteristics by makingchanges to how the aircraft or gas turbine engine are controlled duringtheir use. This may be done, for example, by a control system of theengine (such as the EEC 42) making changes to various control parametersof the engine. Similar changes may be implemented by other controlsystems of the aircraft during use (e.g. during flight). The EEC may bemore generally referred to as an example of a control system 42 arrangedto control operation of the aircraft (e.g. it may be a control module ofa control system).

The present application further provides an aircraft 1 having a fuelcharacteristic determination system 102, 106 according to any one ormore of the examples disclosed or claimed herein. The aircraft 1 furthercomprises a control system arranged to control operation of the aircraftaccording to one or more fuel characteristics determined by the fuelcharacteristics determination system. The control system may comprisethe engine EEC 42, with which the fuel characteristic determinationsystem may be in communication or partly integrated therein. In otherexamples, other control systems of the aircraft may be provided withfuel characteristics and the aircraft controlled accordingly.

The step of operating 1067 the gas turbine engine or aircraft accordingto the one or more fuel characteristics may be performed automaticallyin response to the determination of fuel properties without anyintervention of the pilot. In some examples, it may be performed afterapproval by a pilot, following the pilot being notified of a proposedchange. In some examples, the step 1067 a may include automaticallymaking some changes, and requesting others, depending on the nature ofthe change. In particular, changes which are “transparent” to thepilot—such as internal changes within engine flows which do not affectengine power output and would not be noticed by a pilot—may be madeautomatically, whereas any changes which the pilot would notice may benotified to the pilot (i.e. a notification appearing that the changewill happen unless the pilot directs otherwise) or suggested to thepilot (i.e. the change will not happen without positive input from thepilot). In implementations in which a notification or suggestion isprovided to a pilot, this may be provided on a cockpit display of theaircraft, and/or sent to a separate device such as a portable tablet orother computing device, and/or announced via audible sound such assynthesized speech or recorded message or a particular tone indicativeof the proposed/notified change.

In other examples, the step of operating 1067 the gas turbine engineaccording to the one or more fuel characteristics may include providing1067 b the gas turbine engine with fuel having different characteristicsto that of the fuel for which the one or more fuel characteristics weremeasured in step 1066. This provision of a different fuel may includeloading fuel having different fuel characteristics into the fuel tanksof the aircraft when refuelling the aircraft.

In some embodiments, the one or more fuel characteristics determined mayinclude the density of the fuel calculated from the mass and volumemeasurements. In such an example, the aircraft may be operated accordingto the fuel density.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

1. A method of checking refuelling of an aircraft comprising a gasturbine engine and a fuel tank arranged to provide fuel to the gasturbine engine, the method comprising: receiving an input of calorificvalue data for fuel provided to the aircraft during refuelling;independently determining at least one of: (i) the calorific value offuel supplied to the gas turbine engine while the gas turbine engine isin use; and (ii) the calorific value of the fuel provided to theaircraft during refuelling; and providing an alert when the determinedcalorific value is inconsistent with the calorific value data inputreceived, wherein the independent determination is performed by burningfuel, taken from the fuel tank, in an auxiliary power unit of theaircraft that is not used for propulsive power.
 2. The method of claim1, comprising determining the calorific value of the fuel provided tothe aircraft during refuelling using one or more sensors.
 3. The methodof claim 2, wherein the one or more sensors are provided off-wing at arefuelling site.
 4. The method of claim 1, wherein the determining thecalorific value of the fuel supplied to the gas turbine engine while thegas turbine engine is in use is performed during at least one ofoperations of the aircraft prior to take-off, and climb.
 5. (canceled)6. The method of claim 1, wherein the determining the calorific value ofthe fuel is performed by monitoring engine parameters during a firsttime period of aircraft operation during which the gas turbine engineuses the fuel; and determining the calorific value of the fuel based onthe monitored engine parameters.
 7. The method of claim 1, wherein thedetermining the calorific value of the fuel provided to the aircraftduring refuelling is performed by identifying a tracer in the fuelprovided to the aircraft, and looking up a calorific value correspondingto that tracer.
 8. The method of claim 1, wherein the determining thecalorific value of the fuel provided to the aircraft during refuellingis performed by inferring the calorific value from at least one detectedphysical or chemical property of available fuel.
 9. The method of claim1, wherein the receiving an input of calorific value data for fuelprovided to the aircraft on refuelling comprises receiving data inputvia a user interface of the aircraft.
 10. The method of claim 1, whereinthe receiving an input of calorific value data for fuel provided to theaircraft on refuelling comprises receiving data electronicallycommunicated to the aircraft.